References

Bodies that are optimal at the so-called low critical Mach number M*, at which at least one sonic point on the body flown-over surface occurs, were studied theoretically in Ref. [1]. It was confirmed that M^* achieves its maximum value and, consequently, the wave drag minimum value occurred for the body identical to the Riabouchinsky finite cavity in the classical theory of incompressible fluids. It was experimentally studied in Ref. [12], which demonstrated that in the transonic velocities range the Riabouchinsky half- cavity had the smallest drag among the bodies of rotation with the same aspect ratio  λ=L/D=0.87(where L is the nose part length and D is the diameter of its mid-section). This conclusion is incorrect for aspect ratios λ>2 due to the friction impact the drag as it follows fr om Ref. [24]. The absence of turbulent boundary layer separation from the side surface of the body of rotation under study at zero angle of attack in the range of Mach numbers 0.8≤M_∞≤0.95 was demonstrated in Ref [17].

The main objective of this work is determination of angles of attack α_sep at which turbulent boundary layer separation from the side surface of the studied body occurs. The study was performed with NUMECA FINE/Open software based on Reynolds Averaged Navier-Stokes equations (RANS). The solution of the problem was performed in the framework of fully turbulent flow model without accounting for laminar-turbulent transition using Spalart-Allmaras (SA) and k-ω SST turbulence models. To determine the boundaries of the non-separated flow-around computation was performed in stationary problem setting at various angles of attack. With that, the flow separation indicator was the presence of the zone on the model surface wh ere the friction coefficient C_f < 0. The results obtained with two turbulence models are close to each other, and the difference between the two separation angles does not exceed 1°.

The results of the study obtained for α_sep for the nose part with aspect ratio of are as follows:

α_sep=15° for М_∞=0.5, α_sep=9° for М_∞=0.65,

α_sep=12° for М_∞=0.8, α_sep=13° for М_∞=0.85,

α_sep=5° for М_∞=0.9, α_sep=11° for М_∞=0.95.

Computing results for the longer nose part with aspect ratio are:

α_sep=20° for М_∞=0.5, α_sep=13° for М_∞=0.7, α_sep=21° for М_∞=0.9, α_sep=18° for М_∞=0.95.

The angles of attack α_sep which realize turbulent boundary layer separation from the side surface of the investigated body at free-stream Mach numbers 0.5≤M_∞≤0.95 were obtained. Separation zones location is shown for various models and modes.

The article presents to the development of a technique for buffet initialization boundary evaluation, occurring on a swept height aspect ratio wing at increasing angle of attack during cruising flight modes. The lift coefficient value of the buffet onset is one of the limitations that should be accounted for while designing the win aerodynamic layout of a subsonic aircraft. According to the norms, the margin between the cruise flight mode and the C_Lbuff value or the lift coefficient of the buffet onset should be at least 30%. Thus, knowing the C_Lbuff value through the entire operational speed range is a prerequisite for an aerodynamic wing configuration design beginning from the preliminary design stage. The problem of determining the C_Lbuff has become of special urgency at the transonic speeds due to the substantial aspect ratio increase of the (by 15–20%) of the long-range aircraft wings due to the composite materials application in load-carrying structure.

Despite the successes in CFD aerodynamics gained over the last years, non-stationary separation flow modes are studied, basically, using experimental tools, including wind tunnel tests of airplanes high scale models. Though the cost of such studies is high, they ensure required reliability of the obtained results. It is worth mentioning, that the time consuming computations on multiprocessor computers are costly as well. With this, the high accuracy and reliability of the obtained results are not guaranteed. Preparing mathematical model and building-up computational meshes with hundreds of millions of nodes are commensurable with costs of developing and manufacturing scaled models for tests in the wind tunnel. Thus, numerical methods do not always prove to be less labor consuming and costly than the experimental ones. Despite the fact that computer facilities and software develop rapidly, and the situation gradually changes, experimental methods remain as before the basic tool while performing the studies of complex flows.

The article presents the analysis of typical features of the wing flow-around at the angles of attack corresponding to the start of the buffet mode. The technology of application of the program for computing transonic flow-around based on the full potential method for the buffet initialization computing is demonstrated. Computational results comparison with the data of experimental studies obtained for the model of the wing with fuselage in the wind tunnel is presented.

Vortex formation analysis of the air medium as a gas turbine engine (GTE) propellant allows extracting one of its specifics, namely the gravity character of the technogenic vortices formation. The fudamentals of this vortex formation are being subordinated to the natural vortexes formation of the Earth atmospheric environment. The theory of technogenic vortices of the GTE operating on the ground, the same as the theory of natural atmospheric vortexes is being in the state of its development. It confirms the state of the issues of the working process principle and classification of both technogenic and natural vortices stated in scientific and technical literature and textbooks on the theory of gas turbine engines and metrology. The most informative is the database on meteorological studies of natural vortices (tornados, cyclones, circulations and atmospheric fronts). Thus, due to the technogenic and natural gravity vortex forming similarity, the gist of technogenic vortices' work process should be searched for in the gist of the cyclonic type vortices of the environment. The work process study herewith of the cyclonic type vortices (tornados and others) may be the basis for creating a theory of natural vortex forming.

The problem is set to study the work process of small-sized technogenic vortices with their subsequent adaptation to the work process of natural vortex forming.

The above said problem should be solved relying on the basic equations of gas dynamics (gas flow energy preserving and other) with subsequent yield to the methodology and essence of the gravity type vortices' work process. The most accessible to learning the gravity vortex formation and its vorticity is the energy conservation equation, including its components in the form of internal and kinetic energy of a gas flow, kinetic energy of the environment angular rotation, and heat exchange elements in the form of external mechanic work and heat. Hence, extracting the master unit of vortex formation (angular rotation energy) allows establish functional dependence of the vortex formation under study from the sources of energy capable of generating gravity vortices of various types.

The article presents the methodology of studying, and analysis of the problems of work process cognition of vortex movement of the Earth's ambiences. Classification of vortices according to the gist of their work process. The article indicates the way of splitting the vortex formation into the vortices according to the gist of the work process included into its classification, and further cognition of their physical entity and exploration resulting in vortex characteristics, their consequences and application areas.

By now, several methods for near Earth orbits active cleaning from large-size space debris were suggested. The most difficult stage of such mission is the stage of space debris capture. Capturing method selection and subsequent orbital transportation of space debris depends on its type and angular motion. Rockets' orbital stages may rotate with high angular velocity, which aggravates their capture by manipulators and other means. One of prospective techniques of such object capture is application of a net connected with the space tug by a tether. The object capture by a net can be performed by the net separation with a certain relative speed in relation to the space tug and space debris, or by the net unrolling on the trajectory of the space debris object relative to the space tug. Elastic properties of the net and tether allow reduce the load acting on the space tug while an object capturing process and control the value of this impact.

The paper presents discrete mathematical model of the net movement as a system of material points' elastic interaction, as well as these components interaction with the space debris surface. The possibility of capturing an orbiter type object, rotating with significant angular velocity was demonstrated through this model. The article demonstrates that capturing the object, rotating with angular speed of 5 degrees per second, requires the speed of the net relative to the space debris from 2 to 5 m/s. To capture an object, rotating with angular speed of 30 degrees per second, the net speed should be no less than 10 m/s.

The search for the technologies allowing significantly improve operational characteristics of the prospective civil aircraft was the goal of the work.

The study of innovative technologies, including those, providing the airframe and engine integration, was performed applying methods of alternatives analysis, based on the factorial analysis, and experiments planning assuming performing a series of computing experiments with subsequent comparison of their results.

Three possible innovations trends were considered:

– application of a turbojet engine with the increased bypass ratio for the fuel consumption and noise-at-terrain reduction;

– application of the so-called distributed power plant with the separated gas generator and the fan connected by mechanical transmission;

– airframe and power plant elements integration for obtaining useful effects in aerodynamics and structure, as well as and obtaining new operational properties.

According to these independent principles a number of the long-haul aircraft possible configurations, differing by various combinations of the bypass ratio, the turbojet schemes and technologies of elements integration into the “airframe + engine” system was developed. The number of possible strategies of the integral aircraft, including the base option, corresponds to the number of binomial coefficient of the three factors 23 = 8 according to performing the full-factorial experiment.

A multidisciplinary expert assessment of aircraft configuration options was performed, which turned into the basis for the most effective concepts selection.

Comparison of possible characteristics revealed that some options of airframe and engine integration under consideration had potential for considerable of fuel consumption reduction compared to the conventional long-haul aircraft configuration. The results of the study allow recommend two strategies for further studies. These strategies also possess potential for additional improvement of the other operational characteristics, such as noise-at-terrain and operating safety. Other configurations possess a number of useful elements that can find application while critical technologies development and reduction of technical risks.

The article undertook the attempt to obtain an analytical solution to compute the take-off length of an aircraft equipped with the multi-propeller wing blow-over installation, and estimate the benefits of such engineering decision through the transport operation evaluation of this aircraft.

The main difficulty of this problem lies in the fact that the wing and propeller interaction is an extremely complex and insufficiently studied task. Currently, only approximate semi-empirical formulas for calculating the aerodynamic characteristics of the wing at small relative diameters of the blowing airscrew exist. The exact calculation of the wing flow-around in this case is possible only for strictly specified, nonparametric configurations using the finite-difference method.

In addition, the current complicated situation in the sphere of local air transport in Russia (reduction of airlines and operating airfields) requires the search and evaluation of effective technical solutions for a prospective aircraft of local airlines. One of the ways is envisaged as developing a short takeoff and landing (STOL) aircraft capable of carrying out transport operations in conditions of an underdeveloped airfield infrastructure. It is considered that equipping such a STOL airplane with a multi-propeller electrically powered blow-over system will be an effective solution to the problem. However, the above said complex aerodynamic task does not allow a quick search for the optimal characteristics of this aircraft.

The developed mathematical model, under certain constraints, allows obtain an evaluative analytical solution for the take-off run length of such STOL aircraft, reveal the specifics of parameters interaction and evaluate possible advantages and disadvantages of the aircraft. Within the framework of the model, it was demonstrated that the maximum possible power consumption from the main engines is the optimal value of the corresponding parameter of the electric power plant. The amount of this power consumption determines the blown part of the wing area through the relationship with the critical rotation speed of the auxiliary propellers.

As for performance of a transport STOL aircraft based on L-410, it was shown that a blow-over system based on conventional electrotechnical materials can reduce the take-off run by 30% (up to 300 m), while reducing the payload by 1520% at flight ranges up to 300 km or up to 50% when flying to the maximum range. At the same time, if the electric power plant is designed based on high-temperature superconductors (HTSC), the payload reduction will be much less: negligible at flight distances up to 300 km or about 25% with flight to the maximum range. This allows conclude that the HTSC technology application for such STOL aircraft creation is rather promising.

At present, power electronic components with high heat release have been widely applied in various fields of modern industry. The main problem the developers of the element base are facing consists in creating cooling and thermal stabilization systems capable of removing heat fluxes of high density, while working in a wide range of ambient temperatures. When creating such systems, it is necessary, alongside with the thermal ones, to account for the mass-size characteristics of the device as a whole. Thus, much depends on the heat transfer intensification method selection.

Quite enough attention is paid in modern literature to the issues of radio electronic equipment thermally loaded elements, as evidenced by a significant number of articles and monographs on this topic. Heat release is one of the main causes of the unstable operation of radio electronic equipment. Among the basic factors exerting maximum destructive effect, the increased temperature of the elements is one of the main ones. Thus, the devices operation in the field of the radio electronic equipment is closely associated with heat removal from the thermally loaded components. Depending on the structure and shape of the cooled components, various solutions are employed for organizing continuous heat removal.

Certain problems of large heat fluxes removing in the elements of industrial electronic equipment were considered in [1-3]. The correct choice of the cooling system type ensures trouble-free operation of all cooled components of the device.

A considerable amount of publications in modern scientific publications, both in Russia and abroad, attest to the considerable interest in the issues related to the heat transfer intensification for surfaces of various shapes and sizes as applied to cooling systems for electronic equipment. The issues of heat transfer intensification are set forth in [4-10]. In particular, the criterion equations of various authors for the Nusselt number computing for natural convection are presented in [2, 9]. Experimental studies of the convective heat transfer intensification in rectangular dissected channels and in channels with discrete turbulators were described in [1, 10]. In the studied dissected channels, a process of rational convective heat transfer intensification was implemented, reliably controlled by changing the values of the basic dimensionless geometric parameters. The generalizing dependences for discrete-rough channels were obtained in [7] for free convection conditions, and flow modes and mechanisms of intensification were studied. In [11-14], the authors experimentally studied one of the parameters characterizing the cooling systems both qualitatively and quantitatively, namely, the thermal resistance.

The fins application as a method of heat transfer intensifying leads to the increase in the heat transfer coefficient value by a factor of tens. This method of intensification implies a wide variety of vatious types of fins, such as: longitudinal, transverse, rolling, spiral and many others [15,16]. In [15] the analysis of the expediency of employing different types of fins from the viewpoint of the coolant aggregate state is presented. The optimal edges number selection is presented in [16]. The heat transfer intensification of the systems with a cut-off fin is also considered in [17-20].

The purpose of this work consists in studying the efficiency of the split finning under conditions of natural convection. A test bench was developed for performing the experiment on the study of heat transfer. While the experiments on heat transfer near the cut edge under conditions of natural convection, criterion dependencies were obtained.

Relying on the analysis of literature sources and accounting for the results obtained while experimental studies, the authors established that from the viewpoint of the of axial split finning practical applicability, there are a number of specifics, associated primarily with the fact that the “petals”, obtained as a result of dissection of the heat exchange surface, can be considered as independent ribs. The studies of heat transfer intensification under conditions of natural convection with the cut ribs application were conducted and presented as a result of the work. While the experiment, the effectiveness of the use of split finning is demonstrated, and the most optimal geometric parameters of the working area were revealed. The process of heat transfer was visualized. The boundary layer thickness near the cut edge was computed. Criterion dependencies for heat transfer computing of the systems with axial cut fins were obtained. The prospect of this study is of experimental data verification by numerical modeling programs.

The article deals with the structural damping role analysis while studying the gyroscopic forces impact on the flutter speed. The algorithm for accounting for gyroscopic forces in polynomial Ritz method while computing the aircraft aeroelasticity dynamic characteristics, developed earlier by the authors, was employed. The algorithm was realized in the KC-M software developed in TSAGI and validated while solving aeroelasticity problems in many practical applications.

The computations were performed on the example of the wing of the well-known aeroelastic research model of the four-engine long-distance aircraft EuRAM (European Research Aeroelastic Model), developed and studied in the framework of the European project 3AS (Active Aeroelastic Aircraft Structure). The model characteristic feature is the flutter form occurrence associated with the lateral vibrations of external engines. This form is affected by the gyroscopic forces due to the engines rotating rotors.

The flutter characteristics analysis at various levels of structural damping (characterized by logarithmic vibrations decrement δ ) revealed that the vibrations tones interaction character with account for gyroscopic effect was not principally changed. It was found herewith, that the gyroscopic forces impact on the speed of the considered flutter form might be of different sign depending on the level of the structural damping.

For example, at δ = 0.02 the flutter speed increases by 11.5%, with the maximum value of the engine rotor kinematic momentum (scaled to the model). While increasing the structural damping value, in the beginning, the gyroscopic forces' impact on the flutter speed decreases, it does not practically exist at δ = 0.04, and with further increase of the decrement the impact changes its sign, and the flutter speed decreases. The flutter absence was marked at δ > 0.046 in the range of small rotor rate speed, but while the rate speed increase the flutter may occur. Its speed decreases at that (about 10%) with the rate speed increase. This indicates the importance of accounting for the dynamics of rotor systems while the aircraft aeroelastic phenomena analysis.

The obtained the results were confirmed also by the finite element computing method in NASTRAN system using Rotordynamics module (accounting for the rotor systems dynamics). The computing results on gyroscopic forces impact on aeroelasticity characteristics at various structural damping values performed with KC-M software accord well with computations performed with NASTRAN software.

It was noted that while experimental validation of the gyroscopic impact on the flutter speed of the model in the wind tunnel various results might be obtained depending on the structural damping level. Thus, the detailed computational and experimental analysis of the model dynamic characteristics is required while such tests preparing and running.

The article presents analysis of a concept of the distributed power plant (DDP) while its integration with a “flying wing” type flying vehicle.

A modified airframe model of a prospective long-range aircraft (LRA) of PJSC “Tupolev” development with two power plants integrated into the tail-end was selected as a flying vehicle.

Those power plants represent a bypass turbojet engine where two taken-out fan modules are driven by mechanical transmission from fan turbine of this turbojet. The choice in favor of a mechanical way of power transfer for the aircraft of 2030 level is based on the results earlier performed studies on the engines of new schemes in CIAM named after P.I.Baranov.

The results obtained while numerical modeling of the flow on the upper surface of an LRA airframe were also employed. This modeling revealed that for a long-range flight the mean values of the full pressure's losses prior to the fans differed greatly and depended monotonously on the flow deceleration level in the air intake. According to the calculations, the average value the full pressure restoration coefficients was correspondingly ~0,923 for the first fan module, ~0,952 for the bypass turbojet and ~0,958 for the second fan module.

Refining of the earlier developed model of the distributed power plant was performed to evaluate the impact of the conditions at the inlet of each of fan modules. The performed of mathematical model refinement allowed implementing independent selection of parameters, dimensionality and gear-ratios of reducing gear for DPP fan modules drive, as well as performing independent regulation of output devices of these modules.

The article considers separately the impact of the two main factors on the engine thrust, namely, the fall of the full presure level at the inlet, and its proper heterogeneities.

Calculations revealed that for the earlier selected DPP option while its integration into the flying vehicle under consideration, regulation of nozzles of the turbojet bypass loop and fan modules was required at the cruising mode. With this, gas temperature increase prior to the turbine by ~70 К was required.

Three different variants of the engine which allow excluding the above said regulation were investigated while this work.

The first variant is a version with fans equal by dimensionality and pressure ratio at the designed cruising mode.

The second variant is a version with the first fan module with the pressure ratio increased by 5% relative to the BTJEs fan at the cruising mode.

The third variant is a version with first fan module air consumption decreased by 50% at the cruising mode.

Parametric studies performed employing the develop methodology allowed selecting the degree of bypassing and the degree of pressure increase in the fan optimal by the specific fuel consumption at the cruising mode for each DPP option. The dimensionality of fan modules and main DPP units was refined with account for various losses levels at the inlet.

Analysis of effects associated with the presence of a non-uniform field of the full pressure and leading to its average level decrease at the fan inlet revealed that impact of the presence of non-uniformity might be from 15 to 30% of total impact on the engine thrust.

At the same time, while the power plant parameters selection at the cruising mode with account for the degraded coefficients of the full pressure preservation prior to the fan, the fall of the thrust level due to the proper non-uniformity might be ~2,53% at the given mode. This should be accounted for while selecting an optimal DPP appearance of the configuration under consideration.

At present, a significant part of the research aimed at increasing the energy and mass characteristics of liquid-propellant rocket engines (LPRE) is being performed in the field of new materials application and processing technologies. Other studies are aimed at modernizing the principle of organizing the work process. However, new fuels application is a more long-term and quickly realizable prospect notwithstanding the research costliness in the field of LPRE. Methane may appear one of the propellants, which application may become a new stage in rocket and space industry development. The article considers the historical process of methane formation as a liquid rocket fuel component since it was firstly mentioned by Tsiolkovsky in his works (“Exploration of the World Spaces with Jet Devices”, “Space Rocket”, “Jet Airplane”, “Achieving the Stratosphere”, etc.). A comparative analysis of methane with kerosene was performed in view of the similarity of the work process organization in the LPRE combustion chamber, as well as close hydrocarbon structure. A component close to methane, currently in use in rocket engines, is hydrogen due to the cryogenic nature of both components, which creates difficulties at the design stage of valves, pipelines and gas lines, as well as the organization of the work process in the combustion chamber. Additionally, analysis of the most common fuels based on kerosene, methane and hydrogen was performed. This is especially interesting, since methane fuel pair of oxygen-methane occupies an intermediate position between “oxygen-kerosene” and “oxygen-hydrogen” pairs with respect to the specific impulse and fuel mixture density. The analysis was performed based on physical-chemical, energy, operational, environmental, economic and some other characteristics. This allowed identify the main advantages and disadvantages of methane application as a LPRE fuel and determine its prospects both in Russian and foreign rocket and space industries.

A brief analysis of liquid rocket engines on methane, created or projected in NPO Energomash by V.P. Glushko, KBHA them. S.A. Kosberg, KB Himmash them. A.M. Isaev, the Research Center. M.V. Keldysh, and also to the American firm SpaceX.

Finally, it was concluded that the methane LPRE could replace oxygen-kerosene engines in the near future, since the fuel pair oxygen-methane outperformed the oxygen-kerosene pair by its energy, environmental and economic indicators. Interplanetary flights can become a special field of methane application, since a large amount of methane, the main element of natural gas, can be found almost everywhere in the solar system: on Mars, Titan, Jupiter and many other planets and satellites, which will allow refueling rockets in flight, significantly increasing them the flight range.

The requirements for the ground based gas turbine installations efficiency improvement are constantly increasing.

Heat conversion into work in gas turbine engines, operating by the Brayton cycle, is attended by significant losses, which depend on the cycle parameters and reach up to 60-70% or more. At present, high-tech aircraft engines and their modifications are widely used as ground-based gas turbine plants, making provision for the gas turbines efficiency improvement based on the of combined thermodynamic cycles application.

The article considers the schemes of gas turbine units (GTU) for ground application with combined cycles, allowing improve their efficiency. One of the ways for the gas turbine units cycle improving is heat regeneration of the exhaust gases by installing a heat exchanger at the turbine outlet where a part of the heat is transferred into the air behind the compressor. However, relative bulkiness and substantial weight of the heat exchanger (even of plate type) do not allow at present active application of this scheme in aviation, but it is widely employed in ground applications.

In the case of ground based gas turbine unit, heat exchangers are located in the exhaust chamber or tower behind a power turbine. Thermal ratio of the most widely used tubular heat exchangers is  θ = 0.8-0.9, and plate- type heat exchangers are characterized by the thermal ratio of θ = 0.5-0.8.

It is obvious that the main parameters of the thermodynamic cycle of gas turbines unite, such as the gas temperature T^*g and the compressor pressure ratio (π^* _kΣ ), as well as the parameters determining the working process of additional units (heat regenerators, steam turbine, etc.) of the combined installation play an important role in its efficiency improving. Comprehensive optimization of these cycle parameters is the main goal of the gas-turbine combined unit thermodynamic design.

Computer models of a gas turbine unit with combined thermodynamic cycles developed in the ASTRA SAE-system allowed solve the problems of nonlinear multi-criteria optimization of their operating parameters, determine the most rational schemes depending on the intended purpose and operating conditions of the gas turbine unit.

Russian Turbofan engine TRDDF RD-33 was selected as the basis for studying the heat regeneration impact on efficiency effectiveness. Its low pressure compressor was cut off to eliminate the bypass duct while converting it into ground based installation.

The following variation values of the cycle basic thermodynamic parameters were selected (π*_kΣ = 15, 30, 45, 60 и T^*g = 1200, 1500, 1800, 2100 K). The GTE module without exhaust gases heat regeneration and a GTE with exhaust gases heat regeneration were developed employing ASTRA computer program. The paper presents some results of the study on GTE efficiency improvement.

One of the urgent tasks at present consists in time reduction of flying vehicle launch preparation through introduction of new technologies, equipment and various kinds of tests based on computational experiments.

One of the most expensive types of aircraft is a spacecraft for various missions and tasks in the near-earth space and interplanetary flights. The main system of a spacecraft is the power supply system. Stringent requirements on external impact stability, and operability maintenance in emergency situations, since its failure results in the spacecraft loss. The power supply system preparation and testing is a major part of the testing program and it is performed employing rather labor-intensive methods.

One of the most interesting objects for the test complex optimization are the spacecraft for astronomical observations, as they present a set of a large number of complex technical devices for various purposes and control systems. Such aircraft requires a special approach to ensuring the quality of various electrical systems tests, as well as control and monitoring systems.

It is important to ensure uninterrupted power supply of the onboard service and scientific equipment for the timely data obtaining during the mission. Thus, the task of rapid and high-quality electrical tests performing of such spacecraft is of paramount importance.

The following core systems are being subjected to comprehensive electrical testing: the onboard radio telemetry system, onboard control system, propulsion system, the solar panels orientation system, power system, electrification control system.

Besides scientific equipment, these systems form the basis of almost any spacecraft. Due to the large number of systems subjected to electrical checks, the issue of the electrical tests time reduction, while preserving their quality (guaranteed reliability level of the systems) is relevant. It is necessary to determine the level of reliability and the number of tests based on the system model.

This above said problem can be solved by optimizing the measurement devices' number and functional characteristics, as well as application of automated measurement systems (AIC) for processing a large number of parameters (with account for specifics of electrical tests). This solution allows optimize the testing process, while reducing herewith the number of employed measuring equipment and the test program cost.

The upgraded AIC version allows combine system data into one subsystem of the power bus monitor. It also ensures centralized output of the test data to the Central computer, reducing the number of additional jobs, and simplifying the operators work.

The final layout option has a more optimized build structure. The structural diagram of the modernized automated testing system for electrical testing is presented.

Application of the reliability computation model allowed estimate the number of necessary checks, and the proposed system ensured the electrical tests duration reduction by 30%.

Implementation this complex of has increased the voltage, current and resistance parameters measurements accuracy by 0.01%.

This set allows define the parameters of voltage, as well as the leakage current and the signal occurrence at a specified time instant. This allows localize and fix the problem with the power supply during the test.

The software has a flexible customizable interface that allows quickly respond to changes and emergencies.

The proposed complex can be employed for spacecraft testing and, first of all, telecommunication satellites for various purposes based on of the “Navigator” platform. As examples of products using this platform, we can cite the spacecraft of the Arctic family, Electro and Spectrum.

The article presents the efficiency estimation of an AIST type small spacecraft rotation angular velocity provision and monitoring, employing the improved modification of the software hardware compared to the basic model, undergone flight tests as a part of the flight and trial samples of the “AIST” small spacecraft.

The article deals with magnetometer sensors application as a measuring tool for estimating the rotational motion parameters for various types of spacecraft.

The gist of the article is narrowed down to the fact that for there is no problem of mounting place selection for magnetometers and magnetic actuators (electromagnets) installation on the large-weight spacecraft and space stations. At the same time, for the small spacecraft this problem is topical in view of the impact of onboard scientific and supporting equipment together with magnetic devices of orientation hardware, while magnetic moment work out, on magnetometer sensors measuring data.

The article analyses in detail operation results of a number of devices, containing magnetometer sensors, as a part of various spacecraft types, such as “Foton” No 2,“AIST”, “AIST-2D”. The analysis results revealed that for “Foton” No 2 weighting 6546 kg, almost all deviations in magnetometer sensors measurements could be explained by the measurements instrumental error, while for the “AIST” series spacecraft weighting up to 50 kg the sensors readings demonstrated significant discord.

The authors conclude that the problem of magnetometer sensors' measurement data correctness for small spacecraft is specific due to the essentiality of the impact of scientific and supporting equipment operation, which arises from the dense layout of the payload in the inner space of a small spacecraft.

One of the ways of this problem solving may be the software hardware development, accounting for this impact for each normal operations mode of the equipment.

As an example of such solution, the article presents the software hardware installed on the “AIST-2D” small spacecraft. In the case of the “AIST-2D” small spacecraft, the discrepancies in estimating the angular velocities by two different magnetometer sensors were much lower than for “AIST” small spacecraft. This was achieved by the improvement of software hardware.

The purpose of the article consists in analysis of modern evaluation and modelling methods of contingencies occurrence onboard a spacecraft and substantiation of methods selection for their subsequent application in the process of operative flight control. The process of the aircraft flight control while contingencies occurrence onboard a spacecraft (time limitation for the decision making on contingencies parrying and their type, high dynamics of the processes flow, multicriteriality of the spacecraft current state, the presence of sources of uncertainty) was studied. The inference was drawn, that all considered methods of the onboard contingencies risks evaluation and modelling were not exhaustive. Depending of the current situation, any of the considered methods and on account of the problems being solved could be employed for the contingency risk occurrence onboard a spacecraft evaluation and modelling. The accumulated experience of flight control in manned astronautics revealed that the most interest was provoked by those methods of contingencies risks evaluation and modelling, which reflect the ways of their evolution and aftermath. The selection of the contingencies occurrence risks onboard a spacecraft methods (logic-linguistic and theory of fuzzy sets) is being substantiated by the fact that it allows develop scenarios of the contingencies occurrence onboard a spacecraft and prepare initial data for the decision making on contingencies parrying in case of uncertainty. Methods of contingencies occurrence risks onboard a spacecraft considered in the article may be implemented as tools both in the systems for decision making support on contingencies onboard a spacecraft parrying and in expert systems.

The article deals with the impact of the thermal oxidation process technological parameters, such as the work process temperature and the value of oxidizer reagents consumption, on the oxide film uniformity on the silicon substrate surface. It evaluates the heat-treatment furnace preheating temperature range by modelling in ANSYS with the specified boundary conditions. The homogeneity characteristic, being computed by Min-Max method, was used for homogeneity estimation. The type of silicon wafers under consideration satisfied the number of parameters, such as d = 100 mm, thickness 400 µm, the polished silicon emissivity ε = 0.89, silicon thermal conductivity λ = 31.2 W /(mK), silicon density ρ = 2.33 g / cm³. The preheating temperature range for an atmospheric thermal furnace with a silicon carbide boat is 550–700 °C. Increase in the oxide film distribution homogeneity over the silicon substrates surface due to the uniform distribution of the temperature field was observed as the result of simulation. As a consequence, chemical reactions were close in the intensity of the oxidation processes flow. Experimental validation of the homogeneity increase was obtained due to the thermal furnace boat preheating.

Optimal values of oxidant reagents (O₂, H₂O) consumption at temperatures of 1100, 1000 and 850 °С in the two-phase heating furnace (heating – dry oxidation, work process – wet oxidation) were obtained experimentally. These values allow producing wafers with high-quality silicon oxide (U < 1%).

The article gives technological recommendations on high-quality oxide film provision on the surface of monocrystal silicon wafers by furnace preheating and maintaining temperature between the thermal oxidizing processes, as well as oxidant reagents consumption values selection, allowing producing wafers with high-quality silicon oxide.

This paper presents an innovative energy absorber consisting of an inward folding of composite tube, which is cut axially and turned into the inner of the itself. There is no excess of composite fragments after the composites destruction, and the debris will fill in the inner part of tube to increase the energy absorption. The impact energy is absorbed mainly by the fibers fractures, as well as delamination and friction between composite tube and the cylinder wall of the cap. Impact tests were performed to study the energy absorption performance. To study the shock absorber effect on the shock-resistance of the helicopter crew seat, a four-degree-of-freedom nonlinear biodynamic model corresponding to 50th-percentile male occupant was developed. The simulation results revealed a good shock-absorber shock-resistance performance.

Creation of systems with electric motors application is complicated by the restricted potential for conventional electromechanical transducers characteristics upgrading, namely due to their low specific and voluminous power. In this regard, Russian and foreign scientists are working on creating apparatuses based on high-temperature superconductors (HTS), which, as the studies demonstrate, are capable of ensuring higher values of specific power. In particular, developments of superconductor electric motors for prospective all-electric aircraft, sea vessels electric propulsion and wind-driven power plants are renowned. The article presents the problem solution of computing magnetic fields distribution in the active zone and parameters of excitation and armature windings of electric motor based on high temperature superconductors and ferromagnetic yoke of rotor and stator. The above said problem can be reduced to computing of the magnetic field, created by periodical system of current coils, placed between the two cylindrical ferromagnetic areas under the following conditions: the ferromagnetic sections permeability is assumed infinite, and the motor is considered long enough. The current coils systems herewith may be both of various external and internal radiuses, and of equal ones. As a result, a technique based on Poisson and Laplace equations solution relative to vector magnetic potential was developed. The active zone complicated area of a motor herewith was represented as a set of simple homogeneous partial areas according to the harmonic analysis method. The obtained formulas bear general character and can be employed for the computing parameters of motors of various structural schemes. The proposed technique accounts for the superconductor critical parameters, including the transport current dependence on the external magnetic field; the number of pairs of poles, high-order harmonics impact; the number of slots per pole and q phase. This technique can be applied to the ring armature winding. The solution presented in the article allows determine the HTS motor basic parameters such as quiescent E.M.F., inductive resistance of the armature winding phase, power, weight, etc. HTS generator computing was performed using the obtained formulas. The size of the motor active zone was obtained in accordance with the above mentioned computations. Finite elements analysis of magnetic fields distribution was performed to verify the analytical calculations results and their correction. The obtained results testify the high accuracy of the developed technique.

The proposed unit relates to the field of navigation technology. This type of gyroscopic devices is the simplest and low cost compared to a rotary gyroscope. Single-axis and dual-axis gyroscopes, widely used for movement direction determining, are wellknown, by now, both in civil and military spheres. The main disadvantage of the now-employed gyroscopes is a low speed of the oriented direction determining, reaching up to several minutes. The proposed gyrocompass allows significantly reduce the above said disadvantage and at the same time dramatically improve the accuracy in the “North-South” direction determining. It aims at improving accuracy and speed in the “North-South” orientation determining. This type of gyroscope is much simpler and low cost at commensurable accuracy compared to a rotary gyroscope. The article presents a kinematic scheme of the HS, which allows implement a twin-axis perturbation control method. With this, it allows achieving the required control dynamic characteristics; stabilization accuracy, and the kinematic moment values by selecting the appropriate transfer coefficients Its main property consists in the ability to hold fixed direction of the axis of rotation in space in the absence of the external forces impact on it. This gyroscopic device structure consists of a gyroscope with a rotor of a “brick" shape; communication on angular deviations through amplifiers; torque sensors; and standard stabilizing motors. Thus, the whole complexity of the device consists in the gyro rotor manufacturing, and float chambers in particular. The proposed unit operates in the following way. At the effect of the moment on one of the axes, the rotor gyroscopic moment appears. Pulsating signals along the perpendicular axis appear as well. Thus, there are two types of signals, which can be employed to stabilize the object containing the above said unit is installed. These findings are supported by the presented equations, where  are expressions for the deviations on the two axes. The gyrostabilizer motion was considered in the mathematical model under condition that the values of the moment of inertia on the axes α and β , with account for the presence of additional inserts, would be expressed in the form:  . Thus, the equations were composed with account for these expressions.

The said unit allows find application in the aircraft automatic devices systems.

Scientific results obtained in the work are devoted to the issues of measurement reliability and accuracy of small, slowly changing parameters of low-density gas flows effecting an aircraft in the upper atmosphere. The main trend of the research consists in measuring the aerodynamic forces effecting a flying vehicle in a depleted gaseous environment Measured electric signals coming from primary transducers are often of a low level, sometimes reaching the values less than the noise and interference levels by several orders of magnitude. The most acceptable way of such processes measuring while reaching the specified metrological characteristics is electrochemical modulation (interruption) of the incident gas flow by choppers. It allows obtain the required large gain of the AC signal; easily separate the modulated beam signal from the non-modulated background signal, even if the beam density is lower than this of the background, as well as employ the signal accumulation, which increases the signal-to-noise ratio at the transducer output.

To determine the optimum frequency of electromechanical conversion and ensure the specified accuracy and stability of metrological characteristics, the author proposed employing the method of probabilistic investigation of the stability of the transducer output characteristic. The study of the characteristic's stability is based on the method of probabilistic stability research, ensuring the account for the random character of structural and electro-physical parameters deviations under the impact destabilizing factors and in conditions of mass production. The underlying method of moments allows obtaining the required accuracy with a small amount of computation.

To analyze the stability of the membrane-capacitor transducer transfer characteristic and sel ect the optimum frequency of the flow interruption, programs for computing the nominal values of the frequency response (AFC) and phase-frequency characteristics (PFC) of the transducer, the sensitivity of the response frequency and phase response to the specified parameters were developed. The program for computing the mathematical expectations and dispersion of manufacturing tolerances and temperature coefficients was developed based on the analysis of frequency dependencies of the transducer parameters coefficients effect on its transfer characteristic. Besides, the dependencies of the transducer transfer characteristic fr om temperature at the flow chopper operating frequency, as well as from the temperature change of the transducer membrane were computed.

Based on the measuring transducer transfer function representation in the form of polynomials ratio and analysis of its absolute sensitivity functions to the manufacturing tolerance, structural and electro-physical parameters, as well as destabilizing factors computing dependencies of the transducer transfer function for its stability evaluation and selection of optimal electromechanical conversion frequency were obtained.

The main task of this article consists in improving parallel power active filters operation, with account for the networks with distributed generation specifics, and their operation processes modeling.

Micro-power grids, and the features of their control algorithms, have recently gained considerable attention in a wide range of research community. While the potential for increasing the efficiency, reliability and adaptability of the local grid is the most important motivation for their development, micro-energy systems, in turn, may be implemented to meet the growing demand for electric energy in numerous applications. Compared to the high-power energy systems, micro-energy systems can depend on non-inertial generators, such as photovoltaic batteries arrays, which are connected through inverters. Despite the lack of inertia and other micro-power grids properties, which cause certain difficulties in control, micro-energy systems are controlled better through new control laws, such as, those depending on distributed computations, rather than on centralized processors.

Thus, the distortions of the shapes of the current and voltage waveforms introduced by the high-order harmonic components from the nonlinear load can be supplemented by distortions from the sources of distributed generation and the units for their synchronization with the grid. Wind turbines are the most common renewable energy sources. So the parallel power active filters application is considered in the article for the purpose of compensation of the high-order harmonics, generated in association with their operation features. Currently, there are three most common types of wind generators:

1. Induction generators directly connected to the grid. It is an old concept with a mechanical transmission between the turbine and the generator. The generator operates in a rather narrow range of speeds (just above the synchronous one), the gearbox provides a more or less constant speed of the generator at highly differing wind speeds. The generator requires a considerable amount of reactive power, so often a capacitor bank is connected to it.

2. Synchronous generators with permanent magnets. Usually they are delivered in the kit with rectifiers and inverters. The generator is connected directly to the turbine, and it rotates at a low frequency. The generator is being excitated by magnets, and it is not regulated. The rectifier converts the generator voltage/current into DC. At constant voltage there is a capacitor (for smoothing pulsations and as a certain energy buffer). At the DC voltage side the bulk capacitor is present (for tripples smoothing and as an energy buffer). The DC voltage/current is converted thereafter into AC voltage/current by the inverter with 50 Hz frequency and specified characteristics. Everything is determined by the inverter control system. In fact, the rectifier-inverter is an original DC insertion (the like are employed at the borders between the countries, or for energy transmission over large distances).

3. Double fed induction generators. It is just an induction machine with a phase rotor. The stator of the machine is connected directly to the grid, and the rotor is connected via a rectifier-inverter.

It depends, on many respects, on the concrete case, how separate wind generators are combined into wind farms. However, interference is generated in any case, and an effective compensation system is necessary at each stage.

The parallel active filter circuit and developed mathematical model for the conditions of distributed generation and combined power supply were proposed. They allow efficiently compensate for various harmonic interferences in the grids with distributed generation due to the revealed dependence of the efficiency indices of high-order harmonics correction by the parallel active filter on the value of the supply grid internal resistance and the load node parameters.

Multiphase electric machines with an odd number of phases can become an alternative to three-phase machines in areas where a stable rotation speed within one revolution of the shaft is required, as well as other areas requiring highly reliable electric drives with low noise and vibration, for example in special ventilation systems and complexes .

The problem of formation of the resulting current of 3-phase and 5-phase windings, which are fed from the bridge converter, is considered in the work. A comparative analysis of the resulting currents is given from the point of view of the harmonic composition. In solving the problem, a classical approach was used with the use of the vector method.

It was found that the components of the phase currents of the 5-phase symmetrical winding form the 1^st and 11^th harmonics of the resulting forward current and the 9^th harmonic of the reverse sequence.

The third and seventh harmonics of the phase currents do not form a rotating field; their temporal alternation does not coincide with the spatial alternation of phases. The 5^th harmonic of the resulting current is absent; in the phase-current spectrum of a 5-phase symmetrical winding, the 5^th harmonic component is not contained.

The components of the phase currents of the 3-phase symmetrical winding form the 1^st and 7^th harmonics of the resulting forward current, as well as the 5^th and 11^th harmonic of the reverse sequence.

The 3^rd and 9^th harmonics of the resulting current are absent, because in the phase-current spectrum of the 3-phase symmetrical winding, the 3^rd and 9^th harmonic components are not contained.

By harmonic composition, the resultant current of the 5-phase winding takes precedence over the resultant current by a 3-phase winding. This allows us to assume that within the period of the first harmonic (fundamental rotation frequency) in a 5-phase winding, the vibrations of electromagnetic origin will be less than for the three phase windings.

Experimental studies of prototypes of 5-phase and 3-phase synchronous machines made using identical magnetic systems have shown that the level of mechanical vibrations of a 5-phase machine is lower than that of a 3-phase machine.

Lithium-ion batteries are widely employed for space applications due to a number of advantages compared to other electrochemical systems, especially due to their high specific energy and volume values. Laminated lithium-ion cells possess maximum specific characteristics among lithium-ion cells of other types (cylindrical and prismatic). Besides, they allow use available space more effectively. Despite skeptical attitude towards the laminated Lithium-ion batteries, however, there is information on their successful application in space conditions.

Keldysh Research Center is developing the space oriented Lithium-ion batteries with improved operational and specific characteristics. For example, Keldysh Research Center has already developed and tested the Lithium-ion battery for application in outer space conditions. The objective of the studies being performed is creation of the laminated Lithium-ion battery with improved specific characteristics, large calendar and cyclic life and a possibility to function in conditions of outer space. The task at this stage of the research work consists in searching for cathode and anode compositions allowing achieve maximum specific and volumetric parameters, as well as study electrical characteristics of these batteries under normal climatic conditions.

The studies conducted by Keldysh Research Center with cooperation of the Research Center of Applied Acoustics aiming at searching for cathode and anode composition allowed obtain maximum specific and volumetric parameters for laminated lithium-ion cells (often also called polymer Li-ion by mistake).

Two types of Lithiated oxide of Nickel-Manganese-Cobalt (NMC) and three types of Lithiated oxide of Nickel-Cobalt-Aluminum (NCA) were used as active cathode materials. Six different types of battery-grade carbons were used as active anode materials. N-methylpyrrolidone was used as a solvent for cathode slurries manufacturing, and water was used for anode manufacturing. As the studies revealed, the brand NCA A801-COA was the best choice as an active cathode material and carbon black brand AGP-2A was the best choice as an active anode material. The mass fraction of the active cathode material in the cathode mass was improved up to 93%; the fraction of the active anode material was improved up to 95%. Specific characteristics increase while manufacturing laminated batteries of high capacity. It indicates the batteries manufacturing flexibility while transferring from small-sized laboratory cells to large-sized experimental cells.

The obtained optimal electrodes compositions for Li-ion cells with rated capacity in the range of 1.6-15.3 A·h allow achieving the specific and volumetric values at the level of 230-268 W·h/kg and 520-560 W·h/lrespectively. The following characteristics were demonstrated herewith: for charge-discharge currents of 0.2C-0.2C the cyclability was about 1200 cycles with 97.75% efficiency; for charge-discharge currents of 0.2C-0.1C it was about 980 cycles with 87.8% efficiency, and for charge-discharge currents of 0.5C-1C the cyclability was about 460 cycles with efficiency of 89.1%.

Silumins of the hypereutectic composition in the cast state are characterized by a high level of porosity, the presence of large silicon inclusions and intermetallides, which significantly reduces the range of this material application in industry. To eliminate these drawbacks, samples of hypereutectic silumin (Al- (20-22) wt% Si) were irradiated in vacuum with an intense pulsed electron beam in the surface layer melting mode. Irradiation of the surface of the silumin samples was performed by an intense pulsed electron beam (“SOLO” facility, HCEI SB RAS). The irradiation was performed in a residual argon atmosphere at a pressure of 0.02 Pa with the following parameters: 18 keV; 40 J/cm²; 200 µs; 0.3 c^-1; 20 pulses. The selected mode, as shown by the results of modeling the temperature field formed in the surface layer of silumin, results in the surface layer melting of the material up to 70 μm thickness. Investigations of the elemental and phase composition, the state of silumin defective substructure in the initial state, and after irradiation with an intense pulsed electron beam were performed using scanning electron microscopy (SEM-515 Philips) and transmission electron microscopy (JEM 2100F), X-ray diffraction (XRD 6000, imaging copper-filtered radiation of Cu-K 1, monochromator CM-3121). The samples microhardness was being determined with the PMT-3 device with an indentor load of 0.1 N. The wear-out parameter and friction coefficient were being identified on a TRIBOtechnic tribometer. The results of the studies performed revealed that the high-speed melting and subsequent high-speed crystallization were led to a nonporous surface layer forming of up to 100 μm thickness with the structure of cellular crystallization free of primary inclusions of silicon and intermetallides.

The size of the cells of high-speed crystallization formed by a solid solution based on aluminum was 0.4-0.6 μm. The cells were separated by interlayers enriched with silicon, copper, nickel and iron atoms. The transverse size of the interlayers was up to 100 nm. It was revealed that the nonporous surface layer formation with a multiphase submicro- nanocrystal structure was accompanied by an increase in the silumin microhardness by 4.5 times, and wear resistance by 1.2 times compared to the cast state.

The article suggests combined technique for wings aerodynamic optimization of mini unmanned aerial vehicles (MUAV), which flight modes correspond to critical Reynolds numbers within the range order of 10⁵–10⁶. According to this technique, non-viscous flow-around, flow without separation and aerodynamic characteristics of the finite span wing are being computed in the beginning. The wing planform shape, wing aspect ratio and other geometrics are assumed known and specified. Computation is performed by reliable panel technique. Then the wing profiles shape optimization is performed with account for laminar-turbulent transition, and separation phenomena.

The following assumptions were assumed while wings optimization algorithm developing: the flow-around parameters computation employing 3D analysis model is non-viscous and non-separable. Viscous separated flow-around computing is performed in the contest of 2D-problem of viscous-invicid interaction. Integral aerodynamic characteristics over the wing span are being computed by the technique of lifting line theory using nonlinear section lift data. The suggested technique came from the supposition that aerodynamic characteristics of an isolated wing profile can be extrapolated on the wing. It associates with the fact that the MUAVs wings have, as a rule, a large aspect ratio (AR> 3), and hypothesis of flat sections is applicable for such kind of wings.

The article presents the results of numerical optimization on maximum quality criterion for rectangular wings planform, aspect ratios AR = 5 and AR = 10, at Re = 200 000, as well as arrow-type wing employing the suggested technique.

It was demonstrated that, the moment coefficient constraint allows increase the wing lift-drag ratio, reducing the share of resistance associated with laminar-turbulent transition occurrence and local flow separation formation. At the same time, while optimization in the absence of the moment coefficient constraint each successive quality improvement occurs due to the moment coefficient and wing middle surface curvature increase. The C_ya(a) distribution herewith deviates from the initial one.

The solar battery panels can be divisible by construction into the following types:

– Solar batteries with panels in the form of a frame with the stretched net-like fabric (net-like fabric panel);

– Solar batteries with panels in the form of a frame with orthogonally stretched strings (a string panel);

– Solar batteries with panels in the form of a frame with the stretched flexible film (a film panel);

– Solar batteries in the form of three-layer panels with a honeycomb core (a panel with a honeycomb core);

– Solar batteries with three-layer panels of integral construction (an integral panel).

The structures analysis of various panels reveals that at present all world firms employ generally three-layer panels with honeycomb core.

The structure of such panel consists of carbon fiber-reinforced plastic encasement and a metal honeycomb core.

Pursuing a goal of developing the rigid and light panel, recommendations on selection of carbon fiber-reinforced plastic, honeycomb core, adhesive film and dielectric film are issued based on experiments.

It allowed create lightweight rigid design structure of a solar panel. It was necessary herewith to perform strength, rigidness calculations and vibrations under effect while transportation and operation.

The stress-strain state of panels, forms and natural frequencies were being defined. Calculations were performed by a finite element method in MSC/Nastran.

CQAD4 sheathing element was selected for encasement and honeycomb cores modelling. The CQAD4 element accounts for all internal forcing factors and the encasement geometry, since it perceives membranous, shear, transversal and flexural loadings.

Calculations reveal that tension, occurring in the elements of the offered light-weight structure, have considerable safety margin, and high rigidity at which the maximal shifts do not exceed 0.05 mm, while oscillation frequencies change in within range of 16-91 Hz. The three-layer panel specific mass herewith is only 1.27 kg/m². The structure opens possibilities for further improvement.

The presented work deals with studying the possibility and practicality of high-temperature superconductors (HTSC) application while regional aircraft with hybrid electric power plant (the flying weight of up to 50 tons) developing.

The analysis was performed with mathematical model based on collating the power plant required and disposable power while a cruise flight. The basic energy and mass equations, characterizing hybrid power plants of various structures, including the structure with HTSC application, were derived.

It was revealed, that the turbo-electric aircraft is inferior to the aircraft with conventional power plant in the disposable power value. With application of conventional electrical materials, such as copper, this penalty is significant at any flying weight, and such hybrid aircraft developing is inexpedient. With HTSC technologies application this penalty is significantly lower, but it persists at any flying weight.

It can be explained by the presence of additional units in the power plant, which weight is much higher than the weight of the reducing gear, which they replace. The specific mass characteristics of the units based on conventional electric materials are significantly lower, than for HTSC units, which explains the difference in their application effectiveness. The efficiency change of power transfer herewith is insignificant.

At the same time, it was demonstrated in the framework of the model that the trend of the turbo-electric aircraft upgrading was application of installations and units (both gas turbine engines and electric motors) with the most advantageous specific energy-mass characteristics. With this, as it follows from the derived equations, the power plant should include minimum possible number of electric motors based on HTSC technologies.

It was confirmed in the framework of the constructed mathematical model that if the development of superconductor technologies allows develop HTSC-motors with specific characteristics at the level of 20 kW/kg, then the turbo-electric aircraft disposable power would attain the disposable power values of aircraft with classic power plant. It will ensure unconditional possibility for energy effective regional hybrid aircraft creation.

Ensuring the main rotor blade strength remains as before one of the main problems, the designer faces while developing a helicopter. Heretofore, at the main rotor blade design in a part of static strength ensuring, the designers confined themselves to its computing under the impact of a force of its own weight. While a helicopter operation thereat, the damages of the main and tail rotor units occur after the storm wind impact.

For the main rotor blades, the following situations are possible: the blade spar bending with residual deformation occurrence down to its destruction; corrugation occurrence at the tail sections; the blade contact with ground or the helicopter tail beam. The above said phenomena prove to be possible due to the small inherent rigidity of the rotor blades, which makes them rather sensitive to the wind loading. The designers should take measures on the wind-flow impact protection ensuring while developing rotary wing aircraft.

According to the 29.675 b item of the AC 29-2C recommendation circular, which gives procedures for determining compliance with the requirements of the AP-29 airworthiness standards, when designing the carrier system, it is necessary to avoid overloading the stops and blades in conditions of wind gusts in the parking lot, or from the rotary wing aircraft's main rotor, taxiing nearby.

The article presents a method for computing flexural stresses in an unmoored blade of a helicopter, blown over by a wind flow. It consists in determining the positions of the elastic axis points of the idealized blade model.

These positions fully determine the shape of deformations and, hence, the magnitude of flexural stresses acting in the blade. The initial equation of the blade bending by a wind loading in a linear setting by the Galerkin method is reduced to an equation relative to an unknown deformation coefficient. This coefficient is determined under the condition of neglecting the additional aerodynamic loadings stipulated by the blade elastic deformations, and with their accounting for. The load increase factor was determined from comparing the obtained relations comparison, on which basis the solution allowing avoid the direct integration of the initial equation was obtained.

The equations are presented in a form convenient for numerical determination of the elastic axis points positions of the blade, slope angles and bending moments (stresses). Computation results for the rotor blades of the Mi-8 helicopter are presented. It was shown, that accounting for elasticity introduces significant changes in the bending moments (stresses) distribution along the angle of the blade azimuthal position, which determines the direction of its blow over.

The goal of the presented work consists in developing a technique for computing a part of the aircraft engine casing realized as a thermoelectric generator. The thermoelectric generator (TEG) application onboard an aircraft allows discard the mechanical electric current generator, operating on account of energy extraction from the aircraft engine rotor. At present, a great number of thermoelectric materials, prospective for practical application have been studied. One of prospective trends in this matter is application of housing elements as a basis for TEGs design. The aircraft power plant is undoubtedly the most thermally burdened component. The temperature field along the engine path herewith is characterized by a significant gradient.

Since the thermoelectric gadgets' computing is accompanied by certain difficulties associated with electric and thermal parameters dependencies, the authors developed the technique for computing a housing element, represented in the form of thermoelectric generator of a cylindrical shape. The article presents computation results, performed according to the developed technique, which allow determine and evaluate the value of power output, as well as TEG electric parameters and boundary temperatures of housing walls at the design stage.

The electrical power of the thermo-generator module depends on the flow rate, which cools one side of the housing: a small increase in its speed up to 40 m/s, the power output increases up to 1 kW. It can be seen that under similar conditions with flow rate growth from 50 m/s the power output increases only by 550 W. A similar situation is observed for the case when a TEG is made of of bismuth telluride. Characteristic presented in the article allows determine what engine operation mode would be the most optimal for the TEG effective implementation onboard an aircraft.

To study characteristics and parameters of thermoelectric generator the test bench was employed. The following parameters were measured while the experiments: the resultant current and voltage in thermoelectric modules connected in series (each module is a 64 thermocouples per module, connected in series cased in an insulating ceramic housing), hot and cold junctures temperatures, speed and temperature of the hot and cooling flows.

The paper presents numerical and graphical results of analytical and experimental studies, on which basis the inference can be drawn on the perspective of practical implementation of thermoelectric modules as aircraft engines components. The prospect of TEGs application in high-temperature aircraft and spacecraft power plants is determined by the necessity to obtain powerful enough electric power source onboard with modest weight and size characteristics and high reliability.

Domestic and foreign helicopter building development in the last century opened prospects for convertiplanes application as transportation means for carrying cargoes of considerable weight over the vast territory in conditions of deficiency of the advanced airfield net. Convertiplane is capable of performing vertical take-off and vertical landing. However, convertiplane can ensure significant fuel or electric energy saving while horizontal flight compared to the helicopter of the same weight. Aviation history knows two successful practically realized convertiplane projects. The first project was Canadair CL-84. The second project was the US V-22 Osprey military transport convertiplane. Aero-mechanical schemes of both Canadair CL-84 and V-22 have significant disadvantages. The presented work offers an original convertiplane aero-mechanical scheme, eliminating these disadvantages.

The article lays out the results of studying characteristics of the developed multi-purpose convertiplane, possessing conceptually new aero-mechanical scheme. Various options of the multi-purpose convertiplane, such as ten seats passenger plane, special plane for rescue missions and ambulance, light unmanned convertiplane with high flight duration are considered. Technical characteristics of convertiplane were determined based on the developed technique of preliminary design employing computations of aero-mechanics, aerodynamics, structural strength, weights and centring, as well as comparing the results with the well-known calculation methods.

The results of the studies revealed that among all realization options the offered multi-purpose convertiplane configuration allows achieve higher characteristics, than those of conventional aerial vehicles.The article demonstrates that the existing technical state-of-the-art level allows developing a light multi-purpose convertiplane.

Convertiplane gains its significant advantages through the new turbo-electric power plant, where the last achievements of developing light and powerful electric generators and motors with high power to weight ratio values is employed.

The research paper considers the problem of isochronic curves recovery of thin-walled design structures creep referred to deformations measured within the process of full-scale live experiment. It is known that as time passes creep deformations can appear in the construction. They are graphically usually represented as «deformation-time» curves measured by standardized sample testing. However, it has been found out that the deformation curves obtained during testing procedures of the construction differ fr om standard samples due to various reasons: power-based, technological, thermodynamical, etc. The article presents an approach to the corresponding curves construction, based on the processing of the results of the aircraft construction strength experiment.

Setting up the problem for general thin-walled constructions in mathematical terms, we obtain the necessity to optimize the objective functional in the form of the squared residual error of the corresponded theoretical and experimental deformations to the minimum. Working out the solution of the optimization problem is carried out iteratively using the sensitivity matrix, which is the derivative of the deformation function vector along the vector of elastic parameters variables. As the required parameters which control the stress-strained state (SSS) of the structure we choose the secant elastic modulus of the material. To solve a direct problem of the stress-strained state value determination the finite element method (FEM) in the form of a super-element model is used. This makes it possible to reduce the number of diverse required parameters at sufficient accuracy.

Due to the lack of data from the physical experiment, we obtain the numerical deformation values, using the FEM. This is done by solving a direct problem, wh ere measure of inaccuracy typical for strain and load application gauging is introduced. A mathematical calculation has been made for a four-stiffener wingbox operating under the mechanical and temperature load. Figures of the first and second stiffeners show the change of values of the theoretically obtained deformations in case of iterations in the direction of the corresponding experimental values. Isochronic creep curves have been constructed. The application of the sensitivity function has made it possible to purposefully organize the iteration process in the search for elastic parameters and to construct creep curves for the structural elements. The results of the research can be useful for further development of methods of identifying and improving of thin-walled structures according to the testing data, in case of creeping process as well.

The article suggests several new correlation-regressive models for the aircraft small-sized gas-turbine engines mass computation at the stage of their conceptual design.

A database of the main data and thermodynamical parameters for mass computation of dual-flow turbojet engines was formed. The database includes 92 small-sized turbojets with the thrust less than 50 kN. Equations allowing compute the engine mass at the initial stage of design were derived by correlation-regressive method based on the accumulated statistics.

Model No. 1 uses the total air flow through the engine as an input parameter. The approximating function coefficients were determined based on 88 turbofan engines. The relative standard deviation value for this model was 25.5%.

Model No. 2 uses engine thrust as an input parameter. The approximating function coefficients were determined based on 92 turbofan engines. The the relative root-mean-square deviation value for this model was 18.6%.

The mass model No. 3 uses three input parameters: engine thrust, overall pressure ratio, by-pass ratio. This model involved 77 turbofan engines. The relative root-mean-square deviation value of this model was 13.4%.

The fourth model uses the total air flow, overall pressure ratio, gas temperature in front of the turbine, bypass ratio for calculating the mass.

Statistical coefficients for this model were determined based on 57 turbofan engines. The relative root-mean-square deviation value for this model was 10.1%.

The Kuzmichev mass model depends on five parameters of the gas turbine engine: M_дв = f (m,π_кΣ,G_в,T^*_г, π_в) . The total number of engines used in the statistics was 52. The relative root-mean-square deviation value of this model was 13.5%.

Based on the results obtained, we can draw the following conclusions: at the stage of the gas turbine engine conceptual design, the most preferrable models are model No. 4 and Kuzmichev's model. Models No. 1, No. 2 and No. 3, are most preferable for preliminary estimation of the mass of the propulsion system while an aircraft design.

Recently, the great attention is payed in many countries to the studies aimed at flight cruising speed increase of aircraft of various purposes. The projects aimed at considering the issues creating both passenger (Aerion AS2, QueSST, Sky-lon) and military (SR-72) high-speed aircraft are in full swing abroad.

The similar studies on building-up the flight speed of military planes are carried out in Russia too. Thereupon, the possibility of developing prospective Russian fighter-interceptor based on MiG-31 aircraft, which speed should substantially increased, presents undoubted interest. The same applies to an attack high-speed aircraft of the type of the Soviet T-4 scout bomber, or the US XB-70 “Valkyrie” strategic bomber with maximum flight speed, corresponding Mach number no less than M = 3.

The article presents the results of the study on the power plant optimal shape based on the turbo-ramjet engine with tandem configuration of the high-speed aircraft contours with cruising speed of M_cr = 4.

To solve the stated problem, the software complex consisting of mathematical models of the combined engine, including gas-turbine and direct-flow circuits, supersonic air intake and a full-range jet nozzle, as well as the technique for the aircraft performance characteristic computing. The developed program complex allowed evaluate the efficiency of such combined power plant application as a part of an aircraft with increased cruise speed.

The presented results demonstrated with high obviousness that the effort aimed at the power plant optimal shape formation is most expedient to perform in accordance to the procedure of optimization studies performing, which includes the task setting, the initial data preparation, parametric studies, post-optimization analysis and issuing recommendations.

Parametric optimization with seven parameters and three criteria with goal functions of subsonic and supersonic flight ranges at the optimal altitude, as well as required length of runway for the aborted-continued takeoff, was performed employing the above said approach. The optimization results revealed that the possibility of improving an high-speed aircraft performance relative to the conditionally preliminary basic variant.

Three aircraft options with the highest attractiveness level were selected out of the obtained twenty Pareto-optimal options by the “fuzzy sets” tool. Further final selection of the most expedient one out of these options always up to the development engineer and associated with taking a number of trade-off decisions.

The design of a serial auxiliary power unit was employed as a prototype while developing a new engine. The schematic solution of inlet unit and centrifugal compressor was preserved in the new design, while the engine turbine underwent changes right from the start, since it became radial instead of axial. It required the changes of the combustion chamber design. After studying a number of possible schemes, a decision was made to choose the straight-flow combustion chamber of a ring-type of, which had substantial reserves for minimization on size with relative simplicity of its technological design. The specific feature of this particular combustion chamber is its diagonal positioning relative to the engine axis. A number of problems associated with the lack of experimental and calculation data arise while organizing a working process in the combustion chamber of this type.

The goal of the study is design refinement of the considered combustion chamber structure to optimize the workflow of the annular combustion chamber with the offset zone of a toroid type.

At the first stage, the design refinement of the flame tube structure was performed to organize a vortex structure in the primary zone by changing diameters and a number of clamping apertures and addition of a «springboard» of the internal rim of the flame tube. At the second stage the design refinement of the seat of flame in the primary combustion zone was performed. The atomizer was substituted by the spray injector, and vane swirlers were added to the duct between the deflector and the flame tube wall. The third stage was devoted to the necessary temperature field forming at the combustion chamber outlet. For this purpose the works shaping-up the necessary jets penetration depth, the number and location of shift apertures were performed.

The outcome of the activities consists in obtaining acceptable combustion chamber design of the engine being developed, in which the authors succeeded achieving the flame stabilization in the primary combustion zone, temperature field distribution inside the chamber, excluding its burn-through, and temperature filed irregularity reduction at the outlet.

The article considers solar thermal propulsion (STP) with thermal energy storage (TES) containing high-temperature phase-transition material – beryllium oxide possessing high latent heat of phase transition “fusion-crystallization”. High melting temperature allows obtain the engine specific impulse at a level of 9000 m/s.

Joint optimization of the basic relevant parameters, such as masses ratio of solar mirror concentrator and TES in combination with mirror accuracy parameter was performed. It was demonstrated that that the ratio of TES energy capacity to solar radiation receiver thermal power, or ratio TES energy capacity to solar concentrator area in conjunction with optimal selection of accuracy parameter of the mirror can be accepted as an optimizing parameter. Maximum mass of a spacecraft being placed into geostationary orbit with time limitation of inter-orbital transfer from 30 to 90 days was selected as optimization criterion. Optimization was performed out by Gauss-Seidel method.

The optimization results revealed that optimal ratio of TES energy capacity and light detector power was 22-24 MJ/kW, which corresponds to the optimal ratio TES energy capacity to the concentrator area of 6-7 MJ/m² at rational mirror accuracy parameter of 0.25 degrees. The STP characteristics with TES are presented and analyzed. The article shows that for relatively small flight time of 3040 days optimal values of excess oxidant ratio corresponding to payload mass maximum. The higher value of excess oxidant ratio corresponds herewith to the lower value of the flight time.

Dependences of the TES energy capacity and the concentrator diameter from excess oxidant ratio for a wide interval of flight duration are presented. Expedient areas of heated hydrogen afterburning application for various inter orbital flight duration were determined. The article shows that afterburning is expedient for the time of putting to geostationary orbit of 30 to 45 days. The corresponding excess oxidant ratio changes herewith from 0.3 to 0.1. For the flight above 50 days, the monopropellant hydrogen STP is expedient. Compared to alternative inter-orbit transportation means, employing the combination of small and large thrust engines combination, the gain is about 450 kg under the one and the same inter-orbital transportation time of 60 days.

The article presents computational complex allowing perform computer aided design and calculation of a compact heat exchanger for a small-size gas turbine engine.

The coomputational complex includes a number of blocks based on open commercial programs. The blocks are united by the common software algorithm, developed at Central Institute of Aviation Motors (CIAM).

The input data is changed at each iteration to obtain the required parameters, namely, the regeneration degree and hydraulic resistance.

Computer aided design and calculation include the steps of the initial data entering into the parametric model, checking compliance with the restrictions, automatic model building, meshed models preparation, working medium flowing calculation and computational results output. The initial data is set with account for limitations, such as overall size restrictions and material outlet depth. The possibility of obtaining better thermohydraulic characteristics depending on the model geometry should be accounted for as well.

Automatic building of models is performed according to the set parameters.

At the next stage, the built models are loaded to the ICEM CFD program, and meshes building is performed.

The obtained grid models are used for calculation in Ansys CFX software. Full pressures and temperatures of air and gas at the inlet, as well as the flow rate of gas and air at the outlet are set as boundary conditions. The employed turbulence model is Shear Stress Transport model.

After calculation termination, the resulting file, containing all significant exchanger computational parameters, is formed in the form of a table.

In case of the obtained parameters discrepancy with the claimed requirements, the parameters correction is performed with subsequent repetition of the considered algorithm.

Automation of the design and computing algorithm allows employing it together with CAD complexes for multi-criteria optimization.

The developed computing complex allows obtaining the optimal heat exchanger configuration for a specific task within the specified limits. The calculating complex was being employed in CIAM for the heat exchanger envelope updating, which led to the regeneration degree increase from 62% to 76%, when total hydraulic losses decreased to 1,27% with requirements and restrictions compliance. The genetic algorithm was used as an optimization method.

Lubricating system of the gas turbine engine rotor supports in many ways determines its operation reliability. The experience in developing aircraft GTEs lubricating systems accumulated by many manufacturers is associated with predominant employing of empirical and experimental techniques, and is not practically covered in literature. As a rule, the lubrication systems characteristics are being determined while operating with pure oil, and their recalculation onto the two-phase mixture may lead to overpower of the driving pumps motors and, as consequence, to their weight increase.

The article presents the analysis of the properties of two-phase mixtures pumped through the unit of the aircraft GTE lubrication system based on experimental data. This data was obtained at CIAM with the test bench for semi-natural simulation of lubrication system with oil chamber imitation, and installation for fine-dispersed oil-air mixture forming, where the mixture is formed by the air entrainment effect.

Using the results of oil-air mixture flow visualization, the article shows that in the area of the GTE lubrication systems operating modes the mixture may be considered as one-component homogenous media, possessing the properties of elastic continuum with homogenous sound velocity.

While air entering the exhausting duct the two-component flow of oil-air mixture and air-oil bubbles, clogging the pipe cross section and move relative to the oil-air mixture at low speed is formed.

Characteristics of a discharge gear pump, pumping the oil-air mixture, are affected not only by air and oil properties, but also by the structure formed by the pump throughput capacity.

It has been shown that in GTE lubrication systems a mode of emptying the exhaust gear pump inlet branch may occur with the possible realization of the stratified flow structure, as well as a dynamic locking mode in which a pulsating flow is formed with density waves forming and a polyharmonic fluctuations excitation in the system. Based on the experimental data, the air-oil mixture flow modes map was compiled.

The paper presents the relationships by which give possibility to calculate the thermo-physical properties of the two-phase mixture pumped in the tracts of the GTE lubrication systems. This approach showed good agreement of calculations with experiments in the lubrication system static and transient operation modes.

Turbine disk is the main part of the aircraft engine, since its failure may lead to either emergency or catastrophic situation. According to NTD the GTE load-bearing capacity is being evaluated by the destructive rotation frequency margin, applying the limit equilibrium theory, at destruction along meridian section fr om tangential stress, and at destruction along some cylindrical partially meridian section fr om tangential stress.

Factors affecting the disk load-bearing capacity are the meridian section shape, scheme of destruction (along meridian, cylindrical or mixed sections), the presence of stress concentrators, and the material properties. Allowance for these factors effect on the disk load-bearing capacity while applying the lim it equilibrium theory is not practical.

Destruction of most metals is the result of damages accumulation. Two main mechanism of damages such as voluminous damage (pores growth and merge) and shear damage (cracks growth and merge) are discerned. A model of damages accumulation based on shear damage, i.e. destruction criterion on maximum accumulated plastic deformation, can be employed for numerical determination of the destructive rotation frequency of the turbo-machine disks from nickel alloys.

The plastic flow theory can be employed to determine the disk lim it rotation frequency. A modified version of the classical flow theory with isotropic hardening makes allows implement an arbitrary stress-strain dependence given in the form of strain diagrams.

Several series of calculated overspeed test were performed. The effect of the following factors on the calculated destructive frequency was being studied:

– loading speed;

– the finite elements mesh size.

The computational studies results revealed that the finite element size and mesh computing time did not practically affect the convergence of computation and experiment.

The computational studies results revealed that the finite element size and mesh computing time did not practically affect the convergence of computation and experiment. However, the smaller the grid, the more accurately the cracks development on the disk can be traced.

The obtained computation results were validated based on the results of the overspeed test performed with the low-pressure turbine disk of AL41F-1C engine at the Central Institute of Aviation Motors (CIAM) stand.

The article deals with thermionic cooling system (TCS) of turbine blades (TB) and other hot elements (HE) of aircraft gas turbine engines (GTD), which consists in coating them with a layer (thermionic- protecting layer (TPL)) from heat-proof and heat-resisting material, but with a low electronic work function (EWF). When the TBs and HEs heating, electrons start leaving their surface, taking with them 2-10 MW/m² of thermal energy in exponential-like temperature dependence. It will allow increase significantly the GTE efficiency due to the temperature increase of the working gas prior to the turbine and extra thermionic transformation, as well as increase the GTE reliability and lifespan.

The thermionic cooling technique under development can be employed in aircraft building while creating power gas turbine installations-converters for the spacecraft of increased power capacity and prolonged active lifespan. It can be implemented also while developing commercial systems of putting a payload and tourists into orbit, including a spacecraft based on the reusable first stage of an aircraft type with GTE, or transport aircraft with thermionic GTE. Besides, the technology under development will be called-up for the fuel-and-power sector and shipbuilding while power plants developing, and in oil and gas sector for gas pumping units developing etc.

The TCS realization will allow increase the temperature of the working gas prior to the turbine without increasing the quantity of the air tapped off the compressor, or increase the resource of the most thermally stressed elements of the gas turbine parts, the efficiency increase, thermal stresses reduction in blades due to the thermionic sensitivity to the temperature. It will ensure continuous diagnostics of the turbine state and other high-temperature elements in real-time mode based on electrical engineering parameters, depending on the number of thermo-emission electrons perceived by the anode, and modernize gas turbine installations and GTEs produced in Russia with their resource enhancing due to the extra cooling and without their serious reconstruction.

The article considers the issue of a space vehicles status monitoring automated systems developing.

The goal of the work consists in analysis and application of the tools for flight control systems of conventional spacecraft improving.

The technique of a manned spacecraft current state status monitoring with account for affecting factors (an aircraft orbital movement parameters, structural specifics of an aircraft and ground-based control loop), as well as a throughput of radio communication circuit for telemetric information transmitting, sensor equipment capabilities, onboard measuring instrumentation and computing means were studied.

The conclusion was drawn, that the tasks of controlling processes automation while spacecraft flight control are not exhaustive.

Depending on the designation of an individual spacecraft or spacecraft orbital group not only the tasks and their aggregate set can change, but specific and independent assignments may arise as well.

With account for the current flight control practice in manned astronautics the approach at large to a space vehicles' on-board equipment status monitoring automated systems developing was formed. Automation of problems solving on telemetric information displaying and analysing coupled with information support of a specialist of the group of analysis allows increase the quality level of the managing group functioning. It is achieved through detecting an abnormal situation, potentially translating into emergencies, as well as operational provision of flight control operators with information over a wide range of the problems being solved.

The significance of the spacecraft status monitoring automated systems developing is being proved by the fact that it allows minimize the human factor in the process of a spacecraft control, increase information accessibility and ease-off the burden of analysis group specialist while performing routine operations.

The considered approach to the spacecraft status monitoring automated systems developing can be applied to both the process of of existing manned space vehicles flight control process improvement, and prospective manned spacecraft under development.

The present-day science state-of-the-art allowed ensuring qualitative transfer to many branches of not only scientific but to technologic and other types of human activities. New knowledge aroused at the junction of the well-known scientific and technological trends. Though in certain trends of modern science development some so-called “blind-spots” still exist. The theory of vortex formation is an example of such modern science state-of-the-art. Currently the specialists of this scientific trend cannot establish the physical entity of atmospheric gas dynamic specifics of the vortices under the aircraft air intakes, well as the gist of their work process. The closer analogue of such vortices are the atmospheric whirlwinds, which working processes are associated with the Earth daily rotation. However, the capabilities of modern science do not allow establish the work process of the above said problem gas dynamic phenomena. The scientists in the USA and many other countries declare openly that the do not understand tornado – a small-sized vortex of a cyclonic type. In such circumstances, the scientists are compelled to give definitions to whirlwinds, tornadoes and cyclones by the facts of their physical manifestations in the field of visual perception. Such definitions do not contain the boundary conditions, work process elements, and limit their experimental modeling possibilities. The scientists face a great problem of exploring the work process of the Earth environments vortex forming. One of the main tasks of the Earth environments vortex forming research and its product is establishing the vortex characteristics, their corollary and application areas.

The article discloses the work process of the Earth environments vortex formation. It gives the definition of vortex formation, and specifies the product of vortex formation, including vortex field, tornadoes and air intakes vortices. The work process of vortex formation was established. The article presents the Earth vortex filed characteristics and their corollary.

The Coriolis force role in the process of vortex formation of natural and man-made vortexes was revealed. The results of experimental modeling of vortex formation under the air intake with account for the Coriolis force action are presented.

Nowadays, there is a wide nomenclature of practically new significant tasks of monitoring vast near-Earth space areas by space systems, associated with the space debris problems, spacecraft technical maintenance in orbit etc. All tasks of such kind in an abstract formulation can be interpreted in the form of mathematical problem on optimization of the satellite constellations orbital construction for continuous coverage of specified spherical layers of near-Earth space. However, still there is no theoretical apparatus for effectively solving this problem.

The article formulates for the first time the optimization problems of the two-tier satellite constellations orbital construction for near-Earth spherical layer continuous coverage by the criterion of the characteristic velocity minimum total costs on the system creation. Each tier of such a system is formed in circular orbits with the same altitude and inclination values for all satellites. The satellites of each tier are oriented herewith in such a way that observation cone, formed by the onboard equipment of the satellites in the upper tier are directed downward towards the Earth, while in the upper tier – towards the opposite side.

Decomposition of this problem and its reduction to the traditional problem of selection in the delta-systems class of one-tier orbital constellations and their optimization by the total characteristic velocity minimum was performed in this work. The authors suggest methodological approach to this problem solving; discuss the obtained numerical results and the suggestion on application of the obtained optimal options of the two-tier satellite systems for solving various practical tasks. The two-tier orbital structure in many cases has no advantage over the traditional, single-tiered option. However, under certain conditions the two-tier orbital construction appears after all more preferential.

The presented article deals with the study of technological parameters (temperature and processing time in hexamethyldisilazane (or HMDS) vapours and dehidration) effect on contact angle of silicon substrates pre-processing, including oxidized ones, to evaluate their hydrophobicity. Contact angle measurements were being performed by method suggested by Bickerman. Those angles values were being obtained indirectly due to the known volume and diameter of a water drop. For non-oxidized silicon substrates technological parameters effect on contact angle consists in the following: the 1.5 degrees increase with 30 seconds increase in time of processig by HMDS vapours, 1.2 degrees increase with 120 seconds increase of dehydration time, 0.6 degrees decrease with 45 degrees increase of processing temperature. For oxidized silicon substrates technological parameters effect on contact angle consists in the following: the 2 degrees increase with 30 seconds increase of processing by HMDS vapours, 0,2 degrees increase with 120 seconds increase of dehydration time, 1 degree decrease with 45 degrees increase of processing temperature. Experimental data analysis was performed by Yates analysis, i.e. full fraction analysis. Based on the obtained results the inference was drawn that increasing time of substrates processing in HMDS takes the strongest effect on their contact angle change. Besides, on substrates temperature increase the contact angle decreases irrespectively to the oxide film presence or absence on their surface. The latter, probably, is associated with the fact that hexamethyldisilazane evaporates from the substrate surface, since their maximum heating temperature was close to the HMDS boiling temperature while this study.

The scientific significance of the results outlined in the work consists in studying mechanisms of contact radio interference originating, occurring due to the magnetic field effect of radio transmitters, located onboard an aircraft, on current-conducting mechanical contacts of the structures with non-linear variable resistance, as well as electromagnetic interference generation. The scientific results obtained from the studies demonstrated mechanisms of forming the current, induced by EMF, passing through the aircraft structural elements, which leads to formation of a mechanism of secondary electromagnetic field radiation, interacting with the primary irradiating magnetic field of the radio transmitter. There is a possibility to control the structural elements assembly quality by registering and analyzing the spectral composition of the electromagnetic radiation of the mechanically connected structural elements. Performing complex diagnostics, based on measuring the spectral content of the whole product, and placing antennae around the product under study allows performing reliable estimations of the assembly quality of both separate mechanic components and the entire structure.

The studies performed in this work can be applied to the development and study of a contactless express-method for assessing the structures assembly and erection quality. This method is based on the registration and analysis of artificially generated contact interference under the impact of mechanical vibrations and a high-frequency harmonic electrical signal on the aircraft structures' elements, forming phase-amplitude-modulated oscillation circuits, which can be recorded by either spectrum analyzer or a FAM receiver, or AM oscillations. With this, the levels of their spectral components are measured at a change of mechanical impacts frequency in the range determined by the operating conditions. The measured level of the spectral components of the emitted amplitude-modulated oscillations is compared with the level of the spectral components of the signal emitted by the reference block with given mechanical parameters and normalized level of contact interference.

The main result of these studies allows fruitfully employ the contact interference formation, considered as undesirable phenomenon in the field of electromagnetic compatibility, for estimating the mechanical qualities of the structures (their assembly quality) of various aircraft equipment and units, including assembly and erection quality (especially associated with fixture elements tightening force).

Inertial electromechanical actuators in the form of reaction wheels (RW) found widespread occurrence as actuators of spacecraft attitude control system. The RWs task consists in forming a dynamic controlling moment proportionally to a control signal. This task is reduced to the RW acceleration control.

The article suggests the RW classification, describes advantages and disadvantages peculiar to each of the types. Amid all varieties of the units, reactions wheels with intrinsic disturbing moments' compensation (RWMC) are outlined as one of the most prospective types. The presented work is devoted to the study of these units application possibilities and comparing them with classical ones, where control is performed only by the electromagnetic moment.

To study dynamic and accuracy characteristics of a spacecraft equipped with the RWMC under study, its mathematical model was developed. Analysis of the RWMC dynamic moment development transient was performed. It revealed that transfer functions of compensating and basic (electromagnetic moment control) loops may be represented with high accuracy by the aperiodic link. The time constants of these links were also obtained while the RWMC experimental testing.

The model of the controlled rotational motion accounts for the RWMC static and dynamic imbalances values, as well as the number of RWMC nonlinearities, such as saturation, associated with attaining the limiting angular speed by the rotor and the dead zone while the rotor passes the zero angular speed (for a model without disturbances compensation) etc. Modeling of the spacecraft control system operation in stabilization mode in conditions of ideal measuring of angular position and angular speed was performed to study the effect of the unit specifics on the control system operation.

The spacecraft attitude control system with RWCM was compared to classical RW. In both cases, the control system loop was closed by the PID-regulator, since external disturbances, affecting the stabilization static error value, impact the spacecraft together with disturbing moments.

The simulation results showed that RWCMs has higher accuracy and dynamic characteristics compared to the classical RW. This type of units appears more preferable for developing precise spacecraft attitude control systems, since it allows reduce the “dead zone” of control, as well as oscillation in stabilization transient, especially in the area of near-zero angular rotation speeds of the RW rotor.

Modern power electronics and microprocessor technology state-of-the-art allows develop DC-AC converter with any number of phases in a wide power range.

Realization of a multiphase motor (m > 3) based on the magnetic system of a 3-phase motor is also practically a feasible task with certain modernization of the winding scheme.

As an illustration the article presents a schematic diagram of the four-phase winding, its vector representation, as well as four-phase converter control algorithm while vector pulse-width modulation realization.

The electric drive based on a multiphase motor may display certain advantages in compared to the traditional electric drive based on a three-phase motor and find application wherein the higher requirements are placed on vibrations. The cause of vibrations of electromagnetic origin may be the high-order harmonics of the resulting current, which creates an m.m.f. in the air gap.

Preliminary studies revealed that symmetrical 4-phase winding had the worst figures of the spectral composition of m.m.f., compared to the 5- and 7-phase windings. However, the traction electric drive of the “Granit” electric locomotive was just realized based on the 4-phase asynchronous motor. That is the electric drive based on multiphase motor is already an alternative to the electric drive based on the three-phase motor. It imposes the necessity for comprehensive comparative analysis of multiphase windings and control algorithms for converters to which multi-phase windings are being connected.

The article considers an approach based on classical vector method. With its application harmonic analysis of a resultant current of the symmetrical 4-phase winding. The analysis revealed the phase currents' 1, 5, and 9 harmonics formed the resulting currents of positive-sequence, and the phase currents' 3, 7 and 11 harmonics formed the resulting currents of the negative sequence. Accounting for the fact, that the 1, 3 and 5 harmonics are commensurable in magnitude, significant electromagnetic ripples are theoretically possible within the first harmonic period.

The approach based on the classical vector method considered in the paper can be used to analyze the harmonic composition of the resulting current of multiphase windings with any number of phases. This makes the approach universal for the comparative analysis of multiphase windings on the harmonic composition of the resulting current.

The results of original research on cast structure and properties of a new high-temperature intermetallic gamma alloy Ti-45.5Al-2V-1Nb-1.5Zr(Cr)-Gd-B developed in VIAM [patent RU 2606368] have been discussed. A solidification temperature interval has been determined for the new alloy: solidus temperature 1471°C, liquidus temperature of 1528°C. The pouring gate system has been designed with the help of ProCast software taking into account centrifugal casting technique which provides both full mould filling with molten metal and absence of metallurgical defects in low pressure turbine blade castings. The research was focused on the effect of temperature and duration parameters of centrifugal casting on macro- and microstructures of shaped castings obtained in induction skull ALD Leicomelt 5 furnace. The X-ray spectral microanalysis has revealed that the samples matrix consists of alternating γ -TiAl and α₂-Ti₃Al lamellae; there are areas with lower aluminum content and higher content of vanadium and zirconium/chromium; also excess phases enriched with gadolinium and oxygen have been found (complex gadolinium oxides). Microstructure analysis after hot isostatic pressing has shown that plate-like morphology of structure doesn't change: alternating lamellae of γ and α₂ phases are gathered into colonies within prior β(α)  grains with small amounts of β phase along grain boundaries (the plates possess similar geometrical orientation within each lamellae colony). It has been shown that structure homogeneity of castings strongly depends on pre-heating temperature of casting moulds. As the experiment has revealed the optimal pre-heating temperature of casting moulds for the new alloy falls in the interval 750850°C. The research results have given the opportunity to develop casting and heat treatment processes which allowed to obtain defect-free shaped castings of turbine blades for aviation jet engines.

Cold gas-dynamic spraying is a method for coating process, creation of 3D objects and new materials from powder metals, alloys, composites and powder mixtures. The method was developed based on cognominal physical phenomenon, discovered at the Institute of Theoretical and Applied Mechanics named after S.A. Khristianovich of Siberian branch of the Russian academy of sciences in the early 1980s. Nozzle assembly and a heater are fixed as a part of the cold gas-dynamic spraying test bench based on the industrial robot KR 16-2 in dust-noise proof chamber. While spraying the powder particles are accelerated by the gas flow to the velocities of 400-1200 m/s and form the coating without melting. In a number of works of domestic and foreign researchers the possibility of metallic objects recovery by this method is demonstrated, whereby the study of coatings and materials obtained by this method presents an undoubted scientific and practical interest.

The presented article studies the properties of aluminum coatings formed by the cold gas-dynamic spraying method at corrosion damage zones of the substrates from 1163RDTV structural alloy.

At the first stage of work corrosion damages in the form of surface corrosion of the plates from the 1163RDT alloy were simulated. Then they were recovered by the cold gas-dynamic spraying coatings from ASD-1 aluminum powder. The average measured size of the ASD-1 powder particles was 27 mcm.

Experimental dependencies of porosity and micro-hardness of these coatings and oxygen content in them from deceleration temperature while spraying were obtained. These dependencies allowed sel ect the better coating process mode for performance characteristics recovery of structural elements with corrosion damage.

During the experiments of the second stage the samples recovered by the cold gas-dynamic spraying coatings from the 1163RDTV alloy were tested on tensile strength while static loading. Experimental deformation and fatigue endurance curves were obtained. Due to the low porosity and micro-hardness of the cold gas-dynamic spraying coatings, applied at T₀ = 200°C, the samples with corrosion zones recovered by these coatings were selected for static and fatigue stretching tests. The obtained experimental results analysis revealed that with the considered coating process mode the full static hardness characteristics recovery did not occur. Nonetheless, an A1 recovery by the cold gas-dynamic spraying coating from 1163RDTV alloy increases the sample static hardness characteristics in the elastic region of the deformation curve. The fatigue tests revealed the effect of the stress concentrator on fatigue strength, which should be accounted while cold gas-dynamic spraying application for recovering corroded structural elements.

At the final stage of the work, a coating fr om ASD-1 was formed on the TU-154 stringer fragment (an alloy of B95 series). It demonstrates the ability of applying these coatings on the fuselage frame elements.

The results of the presented work demonstrate the high potential of the cold gas-dynamic spraying method in solving the problems of aircraft construction elements recovery and repair.

At present, designers of almost all foreign countries (USA, UK, France, Germany, Israel, Japan, China and South Africa) decide upon thermal imaging tracking coordinator, employing matrix photo-detecting unit while the type of homing head selection for hew types on missiles. Its modern element base is intrinsically the basis of the fifth generation infrared homing heads. The main advantages of guided missiles of “air-to-air” class equipped with homing head containing matrix photo-detecting unit consist in the presence of significant field of vision, ensuring target patter recognition and its identification, capability of automatic aiming employing and high jamming immunity. All this requires aircraft protection means modernization.

Modern aircraft are being equipped with on-board defense systems, designed to protect an aircraft of various classes and purposes from hitting by aircraft rockets, antiaircraft rocket systems through detecting hazard occurrence and counteracting the attacking means. Onboard defense system “President-S”, “Talisman”, electronic countermeasures equipment of Su-30MKI and aircraft protection system “MANTA” are most up-to-date systems.

The results of performed analysis of modern aviation guided missiles and means of protection from high-accuracy weapons allow conclude that the existing onboard defense systems do not ensure enough level of protection. Namely, they ensure only a passive protection by creating interference action on missiles homing heads, which is inefficient with account for digital signal processing and jamming protection of the guided missiles. Modern heat flares are effective only for protection from the missiles' with single-element photo-detecting unit. Due to target image detection capabilities of modern homing heads with matrix photo-detecting units, the heat flares application is inappropriate. From all the above said, a topical problem of upgrading the onboard defense systems by developing new ways of an aircraft protection from guided missiles follows.

Improving the aircraft protection is possible by active impact on guided missile by protective ammunition included into active protection system, leading to its hitting, self-destruction or mishit.

The goal of the study is enhancing the aircraft protection from guided missiles of “air-to-air” type.

Thus, the developed active protection system is capable of ensuring in automatic mode all aspect detection and tracking of a guided missile, its destruction at a safe distance from the aircraft, in close interaction with the other aircraft systems.

The article presents the results of development and application of the thermal model of a stationary liquid thruster on alternative mono-fuel. It allows calculate the thermal field, and determine internal and external conductive and radiation thermal fluxes, temperature variation gradients speeds in stationary and dynamic modes operating modes of the engine, and calculate heat emission in combustion chamber with subsequent recommendations on upgrading the engine thermal scheme and its reliability.

The purpose of the above said thermal calculation consisted in determining the thermal state parameters and characteristics of the low thrust rocket engine on alternative fuel. The thermal calculations using mathematical model developed and presented in this document solved the following problems: developing the engine thermal model, its verification by the thermal test results, calculation substantiation of the solutions,directed to temperature reduction of propellant delivery valve and capillary delivery tube.

The three-dimensional engine thermal model was built with SolidWorks Flow Simulation 2014, which employs the finite volume method ( a numerical method for integrating the systems of partial differential equations. In heat calculations, the boundary conditions were set identical to the conditions for thermal vacuum tests, which imitated the outer space in full-scale operating conditions.

The experimental data of the engine thermo-vacuum tests, obtained with the development design office Fakel test-bench, were used for the calculation thermal model verification. Verification of the thermalmodel consisted in heaters power selection from the condition of compliance of temperatures in the controlled points and measured ones.

Recommendations on thermal scheme optimization and constructional materials selection were developed according to the thermal calculation results.

Recommendations were also given on optimal structure selection of low-thrust liquid engine on alternative fuel for valve temperature reduction and power consumption reduction while thermocatalytic pack heating-up to +400 °C.

Several design options were considered, and recommendations were given on heat sink application and its impact on the thermal condition of the product, and the effect of the rack material on the thermal condition of the product. According to the results of thermal calculation of the engine structure in functioning mode recommendations are given on substitution of the engine structural elements (heater) and mounting blocks materials not answering the thermal criteria (working values the engine structural units temperature should not exceed qualification value of the temperature).

The article is devoted to the design solutions selection while developing a system of unmanned aerial vehicles in conditions of uncertainty. The presented article such system is assumed as a party of cruise missiles (CM) targeted at hitting an enemy naval ship grouping.

Besides solving the problem of cruise missiles optimal distribution over the target assignments this work solves the problem of ensuring stability at large. Here, the stability means achieving the probability of hitting the targets, no less than the specified one, for all possible values of uncontrolled factors.

By stability in the article is meant the achievement of the probability of defeat of target tasks not lower than given for all possible values of uncontrollable factors. Thus, the problem is set as:

where d is the vector of design parameters, E(ω) is the distribution function, and P is the probability of failure.

The distribution function E(ω) is constructed with engagement of statistical synthesis operations. A regularity criterion was adopted as a criterion of stability:

where Κ_¡  is the Lipschitz constant in the i-th row of the statistical sample of the N volume, Κ_¡pos is the specified value of the Lipschitz constant.

To ensure stable design solution, the contracting mapping is necessary, i.e. the Lipschitz constant should be less than one. With this, the less the Lipschitz constant value, the higher the degree of the design solution stability.

At each step of the statistical sample, two variants of design parameters are set. They are necessary for stability condition calculatiщn. The model values of the Lipschitz constant are restored in the class of trigonometric polynomials:

The problem of CM system optimal ranging is being solved at the already obtained stable vector of the design solution (the set of design parameters) y^ust.

The presented work solved the problem of CM system of optimal ranging, which maintains six target problems. The initial thrust-to-weight ratio and the wing area are assumed as design parameters. The target’s required payload mass, coordinates, speed and course are assumed as uncontrolled parameters.

Three nominal sizes of CMs were considered in the framework of the set problem:

Depending on the uncontrolled factors values, two variants of the cruise missiles optimal ranging were solved, and two distribution functions Ε(ω) were constructed. It is shown that the probability of the system performing the target task appeared to be the same and equals to Ρ – 0,9.

Further, the problem of a design solution selection stable to uncontrolled factors was solved. The stability conditions gave the following design parameters:

Thus, a cruise missile with such parameters solves all the target problems with uncontrolled factors given in the work, i.e. the cruise missile system includes cruise missiles of the same type, and the probability of accomplishing the target problem by the system is 0.9.

For risk reduction of the uncontrolled growth of space debris number at the near Earth orbits, it is necessary to remove the most dangerous objects, such as the worked-out orbital stages of rocket carriers and non-operating spacecraft, which may be the sources of space debris. The most complicated stage of passive uncontrolled object's removal from an orbit is its capturing. Selection of capturing technique is determined by the type of space debris and its angular motion. For example, very often the orbital stages are being purposely spin-up around their transverse axis by jet nozzles to guarantee avoiding collision with the detached payload. It complicates capturing the object of such type by the space tug to remove them from the orbit. While employing manipulators or classical mechanisms of a beam-cone type for capturing, when the engine nozzle of a rocket carrier plays the role of the cone, significant overloads may occur in capturing units.

The presented article proposes employ for docking the expanded beam (strip) with aspect ratio. As in the classical docking technique, the nozzle of an orbital stage is being employed. The docking scheme being suggested allows reduce impact forces occurring while docking with rotating objects.

The docking assembly model was developed to study the effectiveness of the suggested scheme. The flexible beam was modeled by the system of solid bodies (beams) connected by cylindrical hinges. To imitate the bending stiffness a torsional spring was being installed in every hinge. The system model was developed using MSC.ADAMS CAE software. The system model was developed in MSC.ADAMS CAE software.

The process of docking with rotating orbital stage, using the three beams variants of large, medium and low stiffness, was analyzed through the developed model. While docking process, the reaction force value in the hinge, connecting the beam with the space tug hull, maximum tug angular velocity and the success of entire docking operation were controlled. The results of modeling confirmed the impact loads reduction while docking with reduction of the beam bending stiffness. The flexible beam will allow employ greater closing-in velocities with uncontrolled rotating objects of space debris to increase the successful docking probability. The beam elastic properties herewith allow reduce the effect of disturbance forces on the space tug while the beam contact with the docking surface of a docking port (nozzle).

The article is dedicated to the topical issue of helicopter design, namely static stability of an unmoored main rotor blade of a parked helicopter under the wind impact.

The article considers a general case when the wind velocity is directed at an angle to the helicopter longitudinal axis. Velocities and angles of attack of the blocked main rotor's sections are being determined. The authors used experimental data on straight and oblique wings blow-down, as well as circular blow-down of the NACA 23012 profile while determining aerodynamic loads in the blade section, blown by wind flow.

The aerodynamic load, acting in the blade section, is a function of the blade curve, and changes according to the blade's rotation azimuth. Thus, while considering the issue of the blade static stability, the problems on determining the most insecure direction and maximum allowed wind speed of the unmoored main rotor blade under specified position of parked helicopter is solved.

The article considers the blade bending in the plane of least rigidity. The blade torsional deformations are not accounted for while loads determining. It is considered, that the helicopter has main rotor of a common type with hinge mount blades.

Firstly, the solution for the homogenous blade with constant stiffness and aerodynamic characteristics was obtained. The design equation determining the value of wind flow critical velocity in various azimuthal positions was derived. It was established the main rotor blade's stability loss under the wind impact was possible only with oblique blow-down with negative sideslip angles, i.e. when the blade tip position was directed against the wind flow. The wind flow critical velocity minimum value and its corresponding direction were determined. The authors suggest employing the wind coefficient of the blade as a generalized parameter characterizing the blade tendency to the stability loss under wind impact.

Further, the solution for the blade with inhomogeneous parameters was obtained. The value of wind flow critical velocities was obtained by two methods, such as method of straightforward iteration, as well as a method, employing the wind coefficient of the blade.

The article presents the result of the wind flow critical speed computation, performed for MI-8 helicopter main rotor blades, blown-down from the front and back edges.

Perforation of thermal vacuum blanket (TVB) films is performed to ensure vacuum and protection from electrostatic charges effect.

Method of film materials mechanical perforation is the most widely spread for TVB films perforation. With this kind of processing It is impossible to achieve high productivity and perforation accuracy.

Laser perforation of thin materials is one of the high-efficiency technologies for processing materials, and has a number of advantages, such as increasing productivity and perforation accuracy. This method allows quick adjustment of both the diameter, and perforation step.

Fiber repetitively-pulsed laser with the wave-length of 1,062 microns was selected as laser light source. Dot cutting along the hole outline was selected as a cutting scheme.

The process of fiber laser emission action on metalized polyamide films is accompanied by bushy flame in the operation area. The reduction of laser light power and processing speed herewith results in disappearance of bright light emission and significant increase of thermal influence area width up to 300 microns.

From our viewpoint, daisy chain of the following physical effects could serve as such mechanism:

– evaporation of aluminum coating;

– ionization of its vapors;

– impact of this plasma, combined with light power, on polymer, leading to the hole cutting.

One of the evidences of such hole formation mechanism is performed physical-chemical analysis of the obtained holes' edge. The holes edge was studied by electron microscope of JEOL JSM-5910LV series together with INCAENERGY analytic system. The major results of these measurements revealed the carbon content increase in the holes edge area, while oxygen and aluminum content reduced more than three times. Thus, it can be expected that physical process of holes formation with laser perforation of metallized TVB films takes place under combined action of light power and plasma of evaporated aluminum surface layer on polymer base of the film.

At present, the extensive studies of outer space are carried out to obtain scientific, economic and military results.

The solar battery is an important element of a spacecraft since it ensures functioning of its equipment.

The solar battery should have high rigidity at maximum loading factor. The structure rigidity exerts a certain effect on oscillatory process and frequency characteristics while a spacecraft maneuvering. It determines also deformations of a solar battery while its transportation to a specified orbit.

Insufficient rigidity reduces the solar battery efficiency.

The dynamic analysis of solar battery envisages determination of natural oscillations shape and frequency, and a time of the oscillatory process termination.

From these positions, comparison of the two spacecraft “Spectr-R” and 14F150 is being considered.

The finite element models were developed for these occurring while the spacecraft turn along the longitudinal axis were determined.

The inherent characteristics of a solar battery structure were being determined by the finite element method employing “NASSTRAN” software.

To determine values of inherent dynamic characteristics of a solar battery panel a series of simulations of the product dynamics were performed with parameters variation of its mathematical model.

These parameters were determined by elastic and dissipative properties of the solar battery panel.

Comparison of stiffness coefficients values and inertial links damping for these types of spacecraft revealed that the solar panels impact on the dynamic characteristics of these spacecraft was practically the same.

The transient time was of 1000 seconds, which exceeded the admissible values. For the solar battery in the considered configuration, the first mode frequency should be of the order of 0.45 Hz with damping factor of the order of 0.1.

In the considered configuration of the panels, their rigidity characteristics should be 16 times, and dissipative characteristics −3 times greater.

The problem of safety ensuring while an aircraft forced water landing is topical due to periodic incidents while the flights over the water. According to the European Aviation Safety Agency information, the helicopters that have passed the certification procedure topple when performing water landing. This fact indicates that the process of aircraft dynamic contact water surface is insufficiently studied, and the need to account for the aircraft spatial position parameters while water landing.

Confirmation of compliance with requirements of the airworthiness standards (AP, FAR, JAR) while emergency landing and subsequent sailing of the aircraft is based on the results of model tests, which determine the aircraft behavior, the structure loading, its possible destruction and conditions of the most favorable water landing. The hydrodynamic characteristics of models of aircraft fuselages of ground and water basing in the water landing mode, and helicopter equipment with the system of emergency splashdown are studied. A method allowing study such processes at the stage of preliminary design is the finite element method application. However, validity of the results obtained in this way should be verified based on experimental data to enable further practical application of the experience gained.

The article presents the verification results of finite element models of simple geometric bodies (inclined plate, cylinder). These are simplified models that duplicate the shapes of the amphibious aircraft float, the fuselage of the ground airplane and the helicopter ballonet. Verification was perormed employing the concept of Euler-Lagrangian interaction using the generalized “structure-to-fluid” communication simulation algorithm. For the inclined plate, the lifting force coefficients were determined for various deadrise angles and trim at its gliding on incomplete width. A graphic dependence comparing the experimental and computed values was plotted. The changing of overload at the center of gravity was demonstrated for the gliding cylinder, and comparison was performed with experimental data and approximate analytical theory. In all cases satisfactory convergence of the results was obtained.

A helicopter mathematical model with a system of emergency water landing was developed to compute the depth of the ballonet sinking, which determines the level of hydrostatic and hydrodynamic loads. The general case of driving a helicopter to the approaching slope of the wave was simulated with the presence of the initial slip at the moment of contact with the water surface. Based on the graphical dependence of the ballonet transoms movements, a technique for the computed immersion depths determining was formulated. The visualization of the helicopter position change while the water landing process is demonstrated. Based on the developed finite-element model, the other parameters of water landing of a helicopter with emergency splashdown system (overloading in the helicopter center of mass, loads on the fuselage bottom, etc.) can be determined.

The article shows, that a similar approach can be employed to simulate the process of various types of aircraft water landing, including amphibious and ground ones.

Fuel burning in gas turbine engine combustion chamber entails toxic agents formation. Among them, nitrogen oxides and carbonic oxides, which prove deleterious effect upon a human and environment, present the special hazard. In this regard, the article solves the topical problem on upgrading the existing combustion chamber by changing the design of its burner.

At the first stage of the research, several types of burners, differing by nozzle extension geometry, were studied. The studies consisted in determining toxic agents' emissions concentration in the flame formed by the burner.

According to the results of the studies the inference was drawn that the most acceptable burner was the burner with convergent head piece, since it ensures minimum content of nitrogen and carbonic oxides in combustion products. The decision was made on continuing studies of both types of burners, namely, original with diffuser extension and the burner with convergent head-piece, which demonstrated minimum emission of toxic agents.

It was found that the residence time of the burner with converging nozzle extension in the reverse currents zone was 0.15 ms, and 0.025 ms for the burner with convergent head-piece, which is six times less. Testing results were colligated in the form of mathematical dependence of CO and NO from swirl parameter Sg, which characterizes the degree of the nozzle head-piece opening-out.

During the next stage, the studies on determining the throughput capacity of the burners, as well as the quality of air-fuel mixture preparation at their outlet were performed.

According to the results of the studies, it was revealed that due to the high velocity pressure there is no significant jet spreading behind the burner with convergent head-piece. The jet herewith has the high ejection capability and forms narrow flow core, in which intensive fuel and air mixing occurs. The burner with diffusion extension forms a wide concentration field and its low level, which is explained by volumetric recirculation zone.

The combustion chambers tests hereafter on determining thermal field   and obtaining hydraulic characteristics were performed. The measurements showed that at the outlet of the burner with convergent head-piece in the vicinity of thermocouple No 4 the temperature increase was observed compared to the burner variant with diffusion extension. But both cameras ensure temperature field regulated by general requirements.

While next stage the tests of the engines with the combustion chambers under study were performed. The tests data confirmed the reliability of air-fuel mixture ignition during the engine starting. They confirm also correspondence of NK-16ST throttle characteristic to the chambers with both convergent head-piece and diffusion extension in the burner.

The obtained data allowed conclude that employing the burner with convergent head-piece allowed reduce emission of nitrogen oxides by 20% and carbonic oxides by 75%. The main characteristics of the combustion chamber can be affected by changes in the design of the nozzle extension in the burner.

At present methods of blade machines characteristics are widely used by many scientists all over the world. However, the applied methods of problem setting while flow modeling suppose the boundary layer to be fully turbulent in all regions, and do not reflect transient effects actually in the flow in effect. For the flows with low Reynolds numbers the problem setting with no account for laminar-turbulent transition might lead to significant disagreement between experimental and computational results.

The article presents the results of the computational study of Reynolds number effect on the first stage of high pressure K-8B compressor with the low aspect ratio of the rotor wheel blades (RW) (0.729). The stage has the following key geometry and gas-dynamic parameters:

The values of corrected specific mass flow rate through the stage are related to the values at the design point. The compressor stage regulation allows vary the setting angles of the inlet control assembly (ICA) and distributor, though at the rotor rotation frequencies under consideration (100% and 95%) zero angles were set. The ICA row, RW row and distributor row contain 46, 35 and 76 blades respectively. The gaps at the periphery and hub of guiding devices were assumed as 0.4 mm and 0.6 mm correspondingly in the stage model. The rotor row gap was assumed as 0.5 mm. The value of the total temperature at the input boundary condition is 288.15 K. For Reynolds number decrease modeling the values of total pressure were assumed as P_in = P₀, P_in = 0,72P₀, P_in = 0,29P₀, P_in = 0,21P₀, where P₀ = = 101325 Pa is the standard atmosphere. The values of static pressure at the periphery were fixed on the outlet boundary condition.

Simulation of 3D viscous flow in blade channel of the stage was performed with ANSYS CFX SOLVER MANAGER in the setting of 3D averaged Navier-Stokes equations (3D RANS). The computational mesh was created with integrated automatic mesh generator ANSYS TURBOGRID and contains 3643432 elements. The solution for the setting with fully turbulent flow was obtained by Menter SST turbulence model. The calculations accounting for laminar-turbulent transition were also performed. For this purpose the Menter SST turbulence model supplemented with γ − Re_θ transition model by Langtry and Menter was applied. For solutionconcordance, “stage” or in other words “Mixing planes” option was used at the rotor-stator interfaces.

According to the calculation results the stage characteristics degradation between maximum and minimum Reynolds numbers was as follows: adiabatic efficiency η*_ad (4%), pressure ratio ( π* ) at the points of max η*_ad (2.8%), corrected specific air flow rate (1.52%) at rotor rotation frequency n = 100%, and ∆ max η*_ad = 5%, ∆π* = 4.3%, ∆G_cor= 2.3% for n = 95%. Thus, the shift of characteristics corresponding to lower Reynolds numbers occurs to the area of reduced flow of η*_ad and π* . The transitional model addition affects these differences as follows: ∆ max η*_ad= 3.9%, ∆π* =2.2%, ∆G_cor= 1.6% for n = 100% and ∆ max η*_ad =3.7%, ∆π^*=2%, ∆G_cor= 1.6% for n = 95%.

Comparing to the experimental results, obtained for n = 95%, application of transitional model of turbulence increases significantly the accuracy of the numerical study. Namely, deviations between experimental data and calculations with transitional model by values of max η*_ad pressure ratio at the points max η*_ad is less than 1%, while for standard SST model these deviations are of about 2% for maximum Re number, and 3.5% for minimum Re.

Comparing the fields relative to Mach numbers for two models (SST and SST γ − Re_θ ), the basic difference in the flow while laminar-turbulent transition modeling consists in qualitatively true modeling of the processes occurring in the boundary layer. In this case, laminar boundary layer near the front edge of the blades, laminar separation and attachment really exist. Turbulization at the rotor wheel blades occurs at the shock wave location, after which the boundary layer already has turbulent structure for the most part with preservation of a very thin laminar layer. Besides, the changes in flow through the radial clearance in the rotor wheel are being present. For γ − Re_θ “bubble” flow-over while Re number reduction slightly reduces its size. The separation near the back edge herewith becomes more intensive.

The article considers methodical approaches to developing mathematical model using “parallel compressors” method, intended for estimation of inlet flow non-uniformity effect on aircraft engine basic parameters. On the example of a two-shaft turbo-jet engine calculation at two characteristic cruise modes the results of calculated estimation, where the base value of σ_st and averaged value of σ_av stayed invariable, were presented. Parametric calculations herewith were performed for each selected relative value of the reduced pressure area.

It was demonstrated that degree of full pressure inlet non-uniformity effect on the engine thrust at the two considered modes differs significantly. Thus, if at subsonic mode this estimation could be reduced to accounting only for the effect of reduction of the averaged value of the total pressure at the inlet, while at supersonic cruise mode such reduction use might lead to considerable errors. With invariable values of pressure recovery factor at the engine entry, corresponding to the flight speed for the typical air intake, external compression σ_st and averaged value σ_av, the flow non uniformity factor Δσ_nu affects mainly the thrust. The degree of this parameter effect herewith depends significantly on the difference of s_st and s_av.

The obtained results of calculated estimations of temperature field non-uniformity at the engine inlet effect revealed that the dependence of relative thrust reduction only at the cost of relative heating was similar for the two considered modes (transonic maneuvering and supersonic flight). At the transonic mode herewith, corresponding to higher values of the reduced rotation frequency of both compressor stages, the thrust decay occurs less intensely due to relatively smaller decrease of air flow rate through the engine with reduced rotation frequency decrease due to air temperature rise at the inlet. As for the difference between the values of total thrust decay , which does characterize the effect of the input temperature field non-uniformity, with the increase of relative heating at the transonic mode it rises more intensively. It is explained by the fact that at this mode due to the less difference between air consumptions in air-gas channels of «parallel compressors» (more «dence» location of pressure downstream brunches) the speeds difference and, consequently, static pressures between the flows is much greater, than at the supersonic flight mode, which stipulates the higher losses level while these flows interaction.

High specific impulse and low mass-flow rate of ion thrusters (IT) make them increasingly popular choice as a spacecraft propulsion system. Recent missions demonstrate the efficiency of these thrusters in such missions as orbit correction and exploration of Solar system. Moreover, there are many developing ideas of creating spacecraft with IT for wider spectrum of missions. However, IT needs to have a longer operation time, due to the small thrust (about several mN).

As a rule, such thrusters failure occurs due to the destruction of Ion Optics electrodes or failure of electron source. IT needs electron sources as a main cathode (for plasma producing), and as a cathode-neutralizer (to neutralize potentials of ion beam). Hollow cathodes are most used devices for Ion propulsion applications, due to low gas consumption and high electron current density.

Application of lanthanum hexaboride or tungsten with BaO impregnating as an emitter material, leads to the necessity of strict sustenance of hollow cathodes operational parameters. Interaction of emitter material with a small quantity of poison gas leads to its surface contamination and, as a consequence, to decreasing of the recoverable current even down to zero. It leads to more requirements to the gas purity, and hollow cathode handling prior to its placement in space. Moreover, to ensure effective operation, the emitter should be heated up to 0.6-0.8 of its melting temperature by the external heater, which, in turn, causes the emitter material evaporation (life span reduction), power consumption increase and longer cathode start-up procedure.

The problems of high reliability of traditional electron sources for ion thrusters led the authors to the idea apply them as cathode with plasma high-frequency discharge. In such device, plasma is generated and sustained by radiofrequency induction discharge. The absence of “loaded” (high temperature, powerful flows of charged particles) electrodes eliminates all problems of the cathode long-term operation provision. As with hollow cathode, the bulk plasma volume acts as an electron emitter, which allows generate high electron currents. The article describes the scheme of the prototype of this device, and the results of its experimental development. Currents generated by the high-frequency cathode were achieving up to 1.7 A at the input power of 120 W. Effectiveness evaluation of the high-frequency cathode is presented.

The problem of illegal unmanned aerial vehicles of various types detection and identification consists in their low flight heights, small sizes and high maneuverability. The presented article analyses the interrelated optimal selection of detection probability figures in the UAV-Radar system, corresponding to the minimum value of the signal-to-noise ratio at the output of the radar receiving part, i.e. the worst conditions of the UAV detection. The authors suggest a new setting of the problem, associated with several pulses detection at the radar input while the signal-to-noise ratio changing dynamically. The article considers the situation when the detection probability grows with time, and the integral of the sum of detection probability and false alarm probability is equal to a certain constant. In conditions of dynamically changing signal-to-noise ratio with account for the accepted condition of constancy of the integral of the sum while preserving the mutually inverse by nature character changing of detection probability and the alarm probability the problem of optimal interrelation above said probabilities values calculation the is being set. The optimization criterion was formulated in the form of the integral of the well-known expression, determining the interrelation between the signal-to-noise ratio minimum probabilities and false alarm. The gist of the formulated optimization problem consists in finding such probability dependence of false alarm from the detection probability in the series of operations of radar detection with growing detection probability, at which the minimum of the integrated value of minimum signal to noise ratios is reached, ensuring detection of point objects at each radaroperation. Based on the performed analysis the authors obtained the functional relationship of the false alarm probability from the detection probability for scenario, when a pinpoint target in the course of radar detection with growing detection probability is being detected at minimum achievable figure of integrated signal-to-noise ratio at the radar receiver input.

One of major factors affecting the successful UAV performing reconnaissance tasks is the possibility of its coordinates exact localization. The UAV position estimation in such systems is usually performed employing such features as reference points, margins or other images informative elements. Application of such characteristic as reciprocal information for this purpose is also possible. The major shortage of these methods consist in complexity of this function computation in real time scale. The presented article suggests the method, generalizing the main features of localization techniques based on reciprocal information calculation. In contrast to the well-known solutions, localization with the suggested technique is performed based on both information characteristics and optical illuminance characteristics of the analyzed images of various formats. The resemblance of real scene herewith with geo-referenced image is computed by subtracting them from the information characteristics, and for accuracy and reliability of the obtained result, the localization is performed based on multi-format geo-referenced images of the object. The localization problem is solved with this method as a problem of minimization of difference of the total volumes of information, obtained from the real object and reference image in the mode of studying the multi-format frames while meeting some additional condition, specified on total illuminance of the studied and compared images. As applied to the considered problem of the UAV localization, the obtained solution ensures maximum difference of estimations of information volume in the ground scene under study and geo-referenced image. The author concluded that the optimal selection should be considered as such a desired functional dependence, which differs to the greatest extent from the calculated function characterizing the studied extreme localization mode.

Space mission control is an integral part of the control process. It allows obtain a fair presentation on the actual state and functioning of constituent parts of a spacecraft (SC), the degree of tasks implementation and its reaction to control actions.

As a rule, two tasks are solved while controlling. The first one consists in predicting the SC and crew abilities to perform the current flight tasks based on current data, and the second one of no less importance is to detect timely a failure onboard a SC and take measures to its elimination at short notice.

The analysis of the SC onboard systems state adds to the control, but this process is more complicated and it is aimed at revealing cause-and-effect relations of the control parameters both with each other and with external conditions. This analysis is performed for predicting the onboard systems state over the planned flight stages to reveal undesired tendencies in control parameters behavior, as well as for analyzing and revealing the causes of divergences and failures of the onboard system operation. The analysis of the onboard systems states is performed as a rule out of the bounds of a SC operative control loop.

For long-term orbital stations' (LTOS) the analysis of the onboard system state is particularly urgent due to the necessity of ensuring long-term operation in conditions of known restrictions on their structure changing. The flight data generalization and their analysis allow reveal the causes of divergences of the onboard systems states, elaborate recommendations on their elimination of reducing their negative effect, as well as elaborate operational decisions on optimization of the onboard systems operation modes, rational resources consumption, ensuring thereby long-term and effective operation of the LTOS.

The article presents methodological approaches employed for the of onboard systems state analysis with account for collateral data. The operational decisions examples, implemented based on the International Space Station flight data analysis, are considered for the events such as:

– parry the negative impact of the jets of orientation engines of transportation vehicles on solar batteries panels (SP);

– reduce fuel consumption during the SP effectiveness evaluation;

– improving the heat transfer of radiators of the thermal mode provision system during the «solar orbits» periods.

The method suggested in this paper provides estimates for the three projections of wind velocity in earths normal coordinate system using satellite navigation systems (SNS) data, as well as on-board barometric airspeed measurements. The wind speed and its direction are assumed constant for a flight leg of 50-60 s duration. This means, that for the given time interval projections of the wind velocity values on the axis of the normal earth coordinate system are constant. Further, the object and observation models are presented, as well as the identification algorithm accuracy characteristics, obtained from the simulation data processing. The airspeed measuring error effect on the wind velocity estimation is also under discussion. The results, showing the accuracy of wind velocity estimation depending on the constant velocity measurement errors, are presented.

The analysis shows that horizontal projections of wind velocities are estimated with high accuracy (relative errors of 13%), but a certain time interval to obtain the proper degree of identifiability is necessary. After this, the accuracy of estimating the horizontal projections of wind velocities remains at a decent level, and does not depend heavily on the increase of the speed measurement error. The wind vertical projection estimation herewith leaves something to be desired. It makes 3040% even at zero flight speed error, and increases considerably with an increase of speed measuring error. Thus, we may conclude that the suggested method can ensure the good accuracy for estimating the wind velocities along the horizontal coordinate axes, and it is not applicable for estimating the vertical component of wind velocity.

To perform automation of a spacecraft state control, it is necessary to define the scope of tasks, which are most dangerous from the viewpoint of their accuracy of estimate. Intelligent systems application whileoperational flight control does not assume the complete waiving from human in the control loop. It should complement his activities by in-depth and rapid evaluation of a vast amount of information, and help to elaborate the correct reaction to the current state of a spacecraft.

While operational efficiency computation, the nominal time t and its technological delay ∆t , spent for evaluation, is assumed as the main control criterion. This technological delay is associated with the time of data receiving from the spacecraft. The spacecraft normal operation evaluation is important as the main reference point for monitoring of its state changing.

While various operations execution the type of commands issued to the onboard systems to ensure the operation execution, capability of their issuing, as well as the ways of technical evaluation of the state of their execution are accounted for. For control automation, it is necessary also to account for the pre-planned possibilities of organized (nonrandom) effecting affecting its state. From the analysis viewpoint, the flight operation execution switches the spacecraft to a new state. Evaluation of the flight operation and the new state of the spacecraft is the purpose of the flight operation controlling.

The monitoring process includes also performing diagnostics of the spacecraft state. More than one point of its state herewith is determined for the current time (interval). Intelligent system application allows employ all previous diagnostic results and represents the dynamics of the development of the process of changing the technical characteristics of the spacecraft in the past, which can be used to the forecast systematic correcting and increasing its validity.

Operation of the intelligent system in real time mode will allow increase the response rate to anomalies occurrence and their development in time with accurate fixation of the drift of data deviation development. An essential advantage of such systems can be the immunity to accidental failures, such as information loss, as well as the determination of non-obvious changes, which might become a forerunner of failures.

Welding is a complex technological process followed by occurrence of internal residual tensions and deformations of a welded structure.

While producing aircraft It is essential to reduce residual tensions and deformations, since:

– the structure's deformations affect an aircraft external aerodynamic contour reducing its aerodynamic characteristics;

– residual tensions sum up with tensions from external loads on the structure, leading to its destruction;

– residual tensions form volumetric stressed state in separate metal volumes, which complicate plastic deformation of a metal and contributes to its transition to brittle state, leading to local destruction of a structure.

This work is devoted to experimental research on automatic plasma welding application instead of argon-arc welding as one the methods for welding deformations reduction while aircraft structures fabrication.

Plasma welding is the welding performed by directed flow of plasma arc. The plasma arc is characterized by the high temperature (up to 30,000°C), and a wide range of its processing properties. It has much in common with argon-arc welding technology.

The main features that distinguish the plasma arc from the conventional one are:

• a higher temperature;

• a smaller arc diameter;

• cylindrical arc shape (unlike the usual conical shape);

• the arc pressure on a metal is 6-10 times higher conventional one;

• the ability to hold the arc at low currents (0.2-30 A).

Conclusion: the plasma arc is more concentrated, powerful and universal source of heating in compared to the conventional one.

The conducted pilot studies consist in comparing parameters of the samples welded by both automatic argon-arc and automatic plasma welding. Based on the performed work, the following conclusions were drawn:

• the sizes of the weld seam (the width of heat-affected area, the weld seam width, the samples bending angles) made by automatic argon-arc welding exceeded about 1.16 times the sizes of the weld seam made by automatic plasma welding;

• the width of heat-affected area obtained while automatic argon-arc welding exceeded about 1.2 times the one obtained while automatic plasma welding;

• the bending angles of the samples with automatic plasma welding are 2-3 times less than with argon-arc welding.

Based on the above said studies at NAZ “Sokol” the decision was made to implement automatic plasma welding. A new installation for sheets butt-joint automatic argon-arc and plasma welding was developed.

The installation consists of:

• bedplate;

• beams with clamping push buttons and mechanism for converging these beams;

• carriages with plasma gun for automatic plasma welding and a burner for automatic argon-arc welding;

• a mechanism for carriage move along three coordinates: along and transversely to the weld seam axis, as well as up/down;

• supporting devices for sheet billets.

The interface of control panel software is intuitive and provides the following functionality:

1. User identification.

2. Identification of the installation readiness for welding.

3. Welding programs database (DB) creating and editing.

4. The ability of welding the parts of various thickness.

5. Selection the already worked-out and saved welding programs.

6. Control of welding parameters.

7. Logging of the welding process.

The effect of implementing the plasma welding instead argon-arc:

1) Higher labor productivity in view of the higher welding speed (by 3-5 times).

2) Time consumption reduction for products leveling after welding (by 50-70%) due to minimal residual deformations in the weld seam due to more concentrated heating source.

3) Time consumption reduction for welding modes testing (by 50-70%) due to the the stored base of welding programs.

Modern air transportation system is characterized by the great dependence on human, all its elements safe functioning determine the very same “human factor” playing a big role in management and stability of the entire system. In the course of time and aviation industry development, the role of the human factor in aviation accidents is being varied considerably. If for old aircraft, which were difficultly controlled and unreliable the human factor share was 5–7%, in the middle of the middle of the last century it was about 50%, and at present of the human factor is about 80% with the uptrend.

To reduce the risk of an aviation event, the Concept of Crew Resources Managing (CRM), based on the provisions of the human factor, is being actively implemented in modern civil aviation. This is a system of measures aimed at enhancing flight safety and effectiveness by the right implementation of human, technical and information resources, as well improving interaction within the crew, and the crew with the personnel of the other CRM components. CRM is an of practical implementation of the human factor principles.

The human factor as the cause of aviation event implies the human inability to react (interfere with) timely to an evolving or created emergency situation to avoid or minimize of this event aftermath.

One of the most important characteristics of a person is the response time of his reaction. In general, the response time is the time that passes from the moment of the an irritant occurrence to the motional response ending. In civil aviation, the ability to respond to irritants (signals, air traffic controllers' commands, aircraft cabin situations) is instilled in the early stages of training in flight schools. However, the practical development of reactions to events undoubtedly plays an important role in the development of the reaction rate under real flight conditions.

For this purpose, training programs for cadets include tasks for training with a list of events, which are practiced on simulators and imply the occurrence of the cadet's correct response to avoid an emergency situation development.

However, while delivering classes with cadets who are commissioned for a new type of aircraft after flight school graduation, it was noted that in the case of a series of one-type trainings, cadets began foresee a situation that wouldl be set by the instructor and developed while training process.

Thus, the effect of “suddenness” vanishes, and after all, failures or other predicaments, which may occur in flight, cannot be predicted in real flight conditions.

This regularity and foreseeing the possible scenario of situation development is able to abate significantly the pilots skill to respond and solve the unexpected problems and reduce the need for analysis and correct decision-making regarding a particular situation. As a consequence, the pilot's main skill “to fly a few seconds ahead of the aircraft” will be blurred and will subsequently be left without development, which will affect the further safe aircraft operation.

The following experiment was conducted to simulate alike situation with the cadets on the simulator. Ten cadets underwent a total of twenty training sessions, of which in 10 training sessions they knew that a simulated collision with the bird (imitation of broken glass) would be planned, and in 10 other cases the task for training did not indicated the planned collision with a bird.

The results were being recorded as follows: the training number and the number of people who could not properly perform the procedures while collision with a bird (imitation of broken glass) were recorded.

Thanks to uncertainty conditions modeling in virtual space (training device, simulator), these skills can be developed on the ground, preparing the pilot for action in almost any situation, and it does not matter whether the situation is caused by a person, a vehicle, or environment. Skills of action in the face of uncertainty will help the pilots in any case to make right decision and eliminate the problem in time.

Electric motors are one of the main actuating element ensuring aircraft systems functioning. Traditionally, they are employed in fuel pumps, oil pumping pumps, hydraulic stations, automation systems, as fans drive, and wing-flap systems. The variety of problems solved by electric motors on board the aircraft, makes them almost one of the main consumers of electric power.

Currently, several types of electric motors are employed in aircraft fuel pumps, such as DC motors with brush-collector unit, induction motors, inductor and reactive motors, permanent magnet synchronous motors (PMSM) with direct start, and brushless direct current motors (BLDCM). All the listed motors have problems related to energy efficiency and mass and size indicators.

Thus, the main promising motor version for employing in aviation fuel pumps at this stage is the PMSM. A number of scientific and practical works are devoted to the development of the PMSM for aerospace systems. In particular, the specifics of field simulation of the PMSM for aircraft air-conditioning systems and general approaches to PMSM development for aerospace applications are considered. The works are devoted to the study of the PMSM magnetic systems and solving the problems of creating a PMSM control system development. The design features herewith of PMSM for fuel pumps are not disclosed in the literature. Although this type of motors has a number of distinctive features, such as working conditions in the field of low negative temperatures, working capacity at low voltage, employing of graphite bearings, etc. All these specifics do not allow employ the results of the works to the full extent.

Thus, the purpose of this article consists in analyzing the design features of the PMSM for fuel pump by developing and examining the PMSM for fuel pump with concrete geometry with account for real operating conditions and evaluating the prospects for the development of the PMSM for fuel pump.

The article presents the improved version of spacecraft power supply system by rational selection of the hybrid power plant basic elements. Power supply system is the most important onboard system from the viewpoint of energy supply and reliability. Failure of this system entails failure of the whole spacecraft.

The main types of power plants, such as a combination of solar and chemical batteries, installations based on various physical phenomena, and electrodynamic tether systems, as well as nuclear ones are known.

Rational selection of the power-plant basic elements to solve specific problems allows improve technical, mass-and-size and cost characteristics of a spacecraft in total.

The improvement of the power supply system energy efficiency is achieved by special schematic architecture and joint application of chemical and kinetic energy storage devices. The hybrid energy storage device will allow maintain the required energy supply of the onboard equipment and compensate peak energy consumption onboard a spacecraft. This energy storage device includes ionistors. Ionistors serve to compensate fast transients while the installation start-up in orbit. Compensation of the occurring kinetic moment is realized by installing two energy storage devices operating in antiphase. Application of contactless, magnetic, high-temperature super-semiconductor suspension in the flywheel allows significantly reduce mechanical losses and increase the storage time of the stored kinetic energy.

The principle of the above said installation operation in both energy storing mode and energy return to the system to consumers' mode is described. The hybrid energy storage device operation in the process of energy return takes place with rotation speed changing, which leads to the necessity of solving the problem of obtaining the AC of stable frequency at the output. This problem is being solved directly by rotating converter or a specialized inverter. Smoothing the peak loads on the battery by ionistors and the lack of brush gear increase the lifespan of the hybrid energy storage device.

Indicative computations show that application of the hybrid energy storage device allow improve mass-and-size characteristics of the power supply system by 24%. The suggested approach will be employed in further activities associated with enhancing the energy-mass perfection of the spacecraft power supply system.

The article considers methods for frame segment obtaining from the B95 alloy by isothermal forging. This method allows obtaining forgings with minimum allowance for machining and requires value of punching force. Isothermal forging can be a more productive alternative to the now employed cutting operation with NC machine tool. The above said alloy is a certified material for aircraft industry and has a high specific strength. One of the B95 specifics consists in rather narrow deformation temperature region. On the one hand, herewith the forging temperature should be selected as maximum to reduce the required force, and on the other hand, the deformation heating-up may lead to overburning, i.e. irreparable damage of the material, characterized by drastic mechanical properties deterioration. To solve this problem, the authors propose to reduce the deformation loading of the material, which can be ensured by controlling the stress-and-strain state and heating temperature of a workpiece while forging.

The stress-and-strain state of temperature fields analysis was performed with engineering software complex Deform, based on finite element method. Deform software found wide application for the analysis of metals pressure shaping. It allows reduce the design period of the process and cost price, as well as increase the quality of production.

In the presented work several options of isothermal forging of a frame forged piece made of B95 allow were studied with finite element method. While modeling, the initial temperature of the process was being varied, and forging tools of various geometry were employed, as well as the auxiliary operations number. Workpieces of various cross-sections, such as circular, square and rectangular ones were used. The initial workpiece position in the stamp was accounted for. For all cases under consideration, the deformation ratio exceeds the permissible value of 60%, and the process temperature was non-uniformly distributed over the forging cross-section. In a number of cases the conditions that could lead to metal burn-out were observed. It was found, that the most rational scheme is the scheme of isothermal forging, in which a rectilinear pressed rod was used as a billet. Its cross-section area was equal to the section area of the frame forging, and the length of the shelves was 3 mm shorter. This scheme application allows produce forging with equivalent strains of no more than 60%, and allowable deformation heating, which does not lead to the of B95 alloy burnout.

The article presents the results of laminated aluminum-fiberglass composite material, formed by thin layers of aluminum alloy and fiberglass, mechanical characteristics modeling. A modified analytical model of layered material accounting for the presence of metal elastic-plastic layers in the composite structure with bilinear defining relationships is being employed for calculations. For the case of uniaxial tensile, the layer-by-layer analysis of the composite strength is being performed with account for residual tensions formed while the material fabrication. The Tsai-Hill strength criterion was used for fiberglass layers. The moment of yielding commence in metal layers is being determined by Mises criterion. The calculation results determined effective strength characteristics, yield stress and strength limit of composites in conditions of uniaxial tensile. The good agreement of calculation results and experimental data within the 90% of accuracy limits was shown.

The effective Young's modulus of the material in the calculations was 51.5 GPa (49 GPa in the experiment). The apparent yield stress of the composite, associated with the appearance of plasticity in the layers of aluminum, was 230 MPa, which in fact coincides with the experiment. The composite ultimate strength in calculation was 540 MPa (585 MPa in the experiment). In fact, it follows fr om the calculations that the yield stress of metal-polymer composite is determined by aluminum layers yield stress, while the strength limit is determined by the strength lim it of fiberglass layers oriented in the direction of load action. The proposed model allows evaluate the effect of residual tensions on the material mechanical strength characteristics. The results of calculations established that the residual tensions might lead to the composite mechanical properties degradation.

The article studies the possibility of developing technological basics of higher mechanical properties of aluminum casting alloys ensuring, and wrought aluminum alloy while employing aero-thermoacoustic treatment (ATAT).

The share of aluminum allows employed in aviation industry is high. Thus, both casting and wrought alloys find application in aerospace industry. Casting aluminum alloys are used for containers and tanks production. In machine building such casting aluminum alloys as silumin are widely spread.

Aluminum wrought alloys present great interest, due to their higher mechanical properties. They are used for aircraft hulls manufacturing. The above said alloys are employed for manufacturing prefabricated shells of aircraft hulls, representing rigid encasements of rather rigid sheet material, which should resist normal and tangent forces and carry all types of loads.

ATAT employing enables increasing the strength of silumins about 1.4 times, practically with preserving elasticity at the initial level or its slight reduction. Significant holding time reduction was observed as well.

The article studies ways of increasing strength characteristics of extra-high tensile wrought aluminum alloy without significant loss of plastic properties of the material.

The article studies ways of increasing strength characteristics of high-strength wrought aluminum alloy without significant loss of plastic properties of the material.

The ATAT effect on the structure and properties of aluminum casting alloys was revealed, which could be associated with the process of micro-plastic deformation and partial recrystallization while treatment, with diffusion processes acceleration, which ensures grinding of solid solution grains. The redistribution and reduction of macro and microstrains in the material significantly affects its properties.

There are more than 1500 large artificial objects on the near-earth orbits, while only 7% of then are active spacecraft. The remaining objects are space debris. The greatest hazard is presented by the large space debris, such as non-functioning satellites, final stages of rocket carriers, staying on the orbit. Their destruction can lead to grave aftermath, since collision of such object with other the objects and fragments may lead to significant increase of the number of small debris, which, in its turn, can lead to impossibility of safe employing of some near-earth orbits. The space debris removal is one of topical problems, which humanity will have to solve in the nearest future.

A method of space debris removal, and transportation system parameters are determined in many ways by the properties of the garbage being removed. Objects capture and removal by tether systems is one of the prospective methods of large objects, such as non-functional satellites of rocket stages, removal from orbit. The removal of a non-functioning spacecraft with flexible elements herewith is a more complicated task, since the possibility of oscillations of elastic structural elements, such as solar batteries panels should be accounted for, which may lead to their destruction and greater clogging of near earth space.

The article considers cable transportation of a large-sized object of space debris with elastic elements, such as solar batteries. The goal of the work consists in studying the mutual effect of tether oscillations and oscillations of flexible elements while transportation active phase. The article presents the developed mathematical model of the system, consisting of space tug and towed space debris with flexible elements. It considers the simplest case when only a constant thrust force effects the tug. No other forces and moments (such as gravitational) are accounted for.

The transported space debris should not be destroyed while towing, and its attached elements (solar batteries) should not tear away. Otherwise, it may lead to greater clogging of space. To analyze the possibility of destruction and selection of such system parameters that will exclude the space debris structure destruction, mathematical model was developed. By dint of this model, the analytical expressions allowing select the tether rigidity depending on parameters of space debris and mass of the tug were obtained. The article demonstrates the existence of critical tether rigidity, that should be avoided while transportation system parameters forming. Direct numerical integrating of the initial equations of the motion substantiated all analytical and numerical results presented in the article.

The purpose of the study is creating an atomizer for aircraft engines testing while ingress of rain, as well as checking the sample for conformity to the standard requirements for testing facilities adopted for aircraft engines' certification.

The review of fluid spraying problems and methods, which formed the grounds for further selection of the liquid spraying scheme, was performed in this research work. The paper presents the description of the technique for the pursuance of the pilot studies of the atomizer with determination of the parameters such as flow coefficient; water distribution irregularity ratio; rooted angle of a drip stream and drops distribution over the diameter with computation of the average median diameter. It presents also the scheme of installation for complex study of a water drip stream characteristics.

Experimental studies of atomizer prototype models were performed according to the above said technique for the purpose of increase integral parameters of the efficiency, as well as compliance check of range of drops diameters from 0.5·10^-3 to 7·10^-3 m, and the value of average median diameter of 2.66·10^-3.

The tests allowed revealing the relationship between the geometric characteristics of atomizers and drip flows being obtained. Development of the most suitable prototype of atomizer allowed obtain the drops within the certification range with average median diameter of 2,656·10^-3 m.

The results of the work are as follows: the problems of rain imitation were analyzed, the technique for the atomizer testing was developed, and the atomizer design was offered and substantiated. Experimental studies of parameters of the above said atomizer were performed design, and its conformity to certification requirements was confirmed.

The subject of the article “Developing calculation method for forced vibrations of turbomachines of a blisk-type blades” by Shorr B.F., Buyukli T.V., Shorstov V.A., Bortnikov A.D., Salnikov A.V., Frolov V.N. Serebryakov N.N. is the blades of a blisk-type rotor wheels.

The research topic is the effect of amplitude-dependent damping in the material of blades on amplitude of the steady-state resonant vibrations.

The goal of the work is definition of the non-stationary components of the aerodynamic forces and resonant stresses amplitudes in the blades at steady-state vibrations.

The article employs the following assumptions: only the steady-state vibrations amplitudes are being computed. Aeroelastic phenomena relating to blade deformation (both oscillations' excitation and damping) are neglected, i. e. gas exciting forces are defined according to the geometry of air-gas channel elements at a specified operating mode regardless of blade vibrations. Mechanical damping in blades material is amplitude-dependent; i. e. blade behaves as a physically heterogeneous body in the sense of energy dissipation, which heterogeneity depends on variable tensions distribution at each form of vibrations. Damping properties are verified by dynamic tests of samples at various excitation levels and frequencies.

The methodology of the work includes a sequential computational study, which consisting of initial normal modes analysis with definition of the operating mode with possible resonances. It also accounts for of the non-stationary components of the aerodynamic forces definition by solving the Navier-Stokes equation at the operating mode of interest, transferring these components to the nodes to the mechanical finite element model of the blade. Finally, the extraction of the harmonic components of the force, and solving the problem of steady-state vibrations of the blade with amplitude-dependent damping.

Calculations revealed that employing of the constant decrement of oscillations might lead to incorrect results. The difference between calculated amplitudes of the vibratory stresses in the considered example was 25%.

Conclusions were drawn on the method structure, as well as that the considered example of calculating the rotor wheel forced vibrations at resonance with the 13^th harmonic of the flow circumferential irregularity shows the utility of accounting for the dependence of the energy dissipation factor in the material on the vibratory stresses amplitude.

The article presents the results of a pilot study of average heat transfer on a concave surface model with regard to the cooling systems of the leading edge of a gas turbine engines turbine blade with spanwise semi-cylindrical ribs in turbulent flow. Relative curvature parameter was being varied by variation of the momentum thickness. Heat transfer has been studied employing a gradient method based on Fourier-Newton law. A test section was a plane channel of 140 × 100 mm consisting of a straight section and a 90° bend. The concave surface of the channel and the object under consideration had a radius of curvature R_w = 500 mm.

The visualization results revealed that when an undisturbed fluid flowed past the first spanwise rib, the reattachment length behind this rib depended on the surface curvature parameter. The latter is the ratio of the momentum thickness to the surface curvature radius. The increase in this parameter fr om 1.38 · 10^-3 up to 2.5 · 10^-3 resulted in the average of 1.6 times reduction in the reattachment length.

This result derived fro m flow visualization has been satisfactorily confirmed by the distribution of local heat transfer coefficients between the ribs. The reattachment length characterized by the peak heat transfer reduced approximately by 1.4 times. No effect of centrifugal body forces on heat transfer in the flow around the second and third ribs has been observed.

It has been shown that in the case of combined effect of centrifugal body forces and spanwise ribs on heat transfer, these factors do not meet the additivity concept of individual effects due to their mutual coupling. In the considered case, the effect of streamwise curvature of the concave surface was observed only behind the first spanwise rib wh ere the momentum thickness was large. This effect was suppressed further downst ream byboundary layer breakup caused by spanwise ribs. The contribution of centrifugal forces to heat transfer enhancement at a given surface curvature radius can grow if the rib height is decreased while the streamwise rib pitch remains constant.

Purposes and objectives of the article consist in the following: formulating hydrodynamic boundary problem of computation of magnetic-fluid seal (MFS) parameters, which belong to the group of noncontact slot seals operating as a hydraulic lock. While developing MFS the annular packets of conducting magnetic brushes were used as magnetic field concentrators instead of teeth. A magnetic fluid resides between the bristles of these brushes in a narrow annular channel. Such a seal gives the minimum friction between the interfaced parts. Numerous calculation methods for the abovementioned hydrodynamic boundary problems developed historically almost independently from each other. General principles for creating numerical methods acceptable for all hydrodynamic boundary problems in general were slated. The aggregate of these concepts and methods allows eventually reduce the algorithm for solving complex hydrodynamics boundary problems to algorithms for solving simple problems of standard structure. An integral relations method employed in this work was developed based on conservation laws and eventually reduced to ordinary differential equation solving. At the domain boundary herewith the boundary conditions are specified both at the rigid wall and the boundaries through which the flow inflows and outflows. Additionally, adhesion conditions are specified. The presented article formulates the new boundary conditions of tracking-concatenation of viscous incompressible flow for internal flows in narrow annular channels. It gives mathematical formulation of the boundary problem for viscous incompressible magnetic flow with possible internal backflows, which detection experimentally is impossible. The boundary problem was set and algorithm for computation of viscous magnetic liquid flow field in the annulus with movable walls of the magnetic-fluid seal (MFS) by the structured method with the exact fulfillment of the boundary and initial conditions was presented.

The article shows that application of mathematical apparatus for solving the boundary problems by the structured method allows calculate in total parameters of the magnetic liquid flow: heat flows, coefficients of friction, heat transfer and distribution of these parameters through the radial clearance of annulus, revealing the areas of potential backflows.

The results of this work may be useful while developing and computing new type of magnetic-fluid seals (MFS) for high-speed shafts of structures and units for various industrial purposes.

The article presents the analysis of a distributed power plant concept for perspective long haul passenger aircraft, which is intended for ensuring more deep integration of a power plant and a flying vehicle, as well as enhancing its fuel efficiency.

While employing an aircraft engine of such kind, separate modules of a power plant may be installed both in the engine nacelle and inside an airplane fuselage, made according to a “flying wing” scheme.

A portion of a boundary layer, formed at the surface of an aircraft, gets into the inlet plane of fan modules, located at the top surface of the fuselage.

The variant of a submerged engine inside an aircraft assumes the presence of a rather long curvilinear intake channel, in which local separations and vortexes inevitably occur. It leads to additional losses of full pressure at the engine inlet.

The article considers separately the effect of two main factors on the engine thrust, namely, the drop of overall level of the total pressure at the engine inlet and its non-uniformity.

To evaluate the effect of the above said components, the results of preliminary work out of the distributed power plant parameters, obtained at CIAM, named for Baranov, in the activities progress on the engines' schemes of new types, were applied.

Calculations were performed employing the first level model of an aircraft gas turbine engine.

Parametrical studies performed using the developed technique allowed select an optimal degree of double-flowness on specific fuel consumption at course speed, and the degree of pressure increase in the fan. The fan modules' and main engine components dimensionality was redetermined with account for various losses levels at the inlet.

The effect of engine parameters changing on the its mass estimation value was performed with the developed modular technique, based on the idea of impeller machine mass proportionality to compression specific work and corrected specific air consumption. The modular technique coefficients characterizing the weight fraction of the turbojet modules were determined based on estimations obtained for detailed element-by-element mathematical model of mass, in the activities progress on the engines' schemes of new types, at CIAM, named for Baranov.

The obtained results of the parametrical studies make it clear that on deterioration of the factor of total pressure preservation at the inlet by 2%, minimum specific fuel consumption at a cruising mode would be achieved in the distributed power plant with double-flowness reduced by 3%, and the total pressure increase degree in the fan reduced by 0.6%. At the same time specific fuel consumption increases on 6-7 % of percent. The specific fuel consumption herewith is increases by 6-7%.

The power plant weight, without account for the weight of the remote fan modules transmission drive may increase by approximately 4-5 %.

Analysis of the effects associated with the presence of non-uniform total pressure field, resulting in its averaged level reduction at the fan inlet, revealed that the effect of non-uniformity presence itself might be of 15 to 30% of the total effect on the engine thrust. It should be accounted for selection of the distributed power plant shape of the configuration under consideration.

Foreign-made aircraft (AC) ingress to domestic civil aviation airlines revealed a number of significant challenges, including the testability provision (abbreviated Tst). We will denote testability management hereafter by FTsbl symbol with Tstblt subscript. An ill-considered implementation of foreign-made components in aviation transportation system of Russia without comprehensive accounting for operation and maintenance factors leads to above-level downtime of cost intensive aerotechnics, and upset of calculated value of an aircraft testability. At present, revealing assessments and factors of testability management, gains special topicality and requires comprehensive analysis. Many scientists in Russia, including V.S. Shapkin, N. Gipich, .G. Evdokimov, A. Stepanov, V. Viktorova and abroad, including Douglas, T. Ross, studied testability as the means of equipment failure-free operation provision through its whole life cycle. However, the studies of testability provision while operation are insufficient. The testability management system is being reduced to compliance with the State Standard 27518-87 “Products diagnosis”, i. e. to totality of coordinating activities on management state, as a part of general enterprise management. These activities are not oriented with respect to testability while operation. They are fulfilled without adequate theoretical development on substantiating the required acceptable testability level of object under operation and control action. It does not achieve the desired goal since functional dependencies of testability management, controlled parameters and acceptable limits of testability parameters variation of controlled products are not substantiated theoretically.

The objective of this article consists in studying the possibility of testability management while operation and developing mathematical model of testability management of an object on the example of testability management of aerotechnics.

The article describes the testability as a function of the monitored object under operation. It presents description of testability computation models and algorithms. Based on the theory of optimal processes and Pontryagin's maximum principle the mathematical model of the function test was studied. A mathematical model of an operated object testability management on the example of aerotechnics. This model is based on measuring indices and parameters of operation, processing of the obtained data, analyzing and developing control action on the operated object.

A mathematical model of controlled object under operation testability on the example of aviation technology, based on the measurement of parameters and operating parameters, the processing of this data analysis and generation of control action on the object of exploitation. An approach to testability management of an object under operation was deduced.

The main modern aviation development trends are based on the fact that aircraft qualities are defined not only by carrier characteristics but also by onboard equipment complex capabilities.

High rates of airborne equipment development and creation intrinsic to recent years have come into contradiction with long service life of airframe and engine. Resolution of this conflict supposes abruption of an aircraft common life cycle as an aggregate of an aircraft and its equipment and shift to logically interrelated separate cycles of aircraft and onboard equipment complex development.

The problem discussed herein consists in optimum employment of cost and time resources in aircraft flight and engineering test practice.

Particularly, the flight and engineering tests are described, which essence consists in giving the answer to the question on how the flight task was realized with the accuracy not worse than the specified one.

Parametric expected uncertainty within the problem formulated has some specific distinctions from situations discussed within statistical decision theory.

First of all, the values of the parameters, which define the hypotheses under checking. These parameters, H₀ and H₁, are not defined (a priori) within the sets of their values Ω₀ and Ω₁, responsible for the system state (H₀ – the system complies with the requirements, H_{1 –} it does not), and are defined in the sense of the system state (“YES” – H₀, “NO” – H₁).

Secondly, inasmuch as on the assumption of employing information methods for optimization of surveillance planning stage at the interval of an aircraft's ground tests the situation, when the probability in the context of alpha and beta errors is required, is inadmissible. The decision making in this case will turn out to be unobtainable due to the lack of information in the sample of observations.

Substantiated information and cost approach, general formulation and the ways of resolving the problem of surveillance of ground measuring complex means while performing aircraft flight and engineering tests, ensures the effectiveness of flight tests with existing test pattern and requirement for minimum consumption of all kinds of resources.

Proved relationship and interpretation of the results open a possibility of obtaining analytical expression of informational measures necessary within the framework of the problem discussed and formulation of the task for ground measuring system equipment observation plan optimization.

While application of this method, the effectiveness of proposed models was about 9 –15 % of augmentation in terms of economic indicators, and instruments and general structures controllability by 15 – 20 %. Thus, general effectiveness of the proposed model equals to about 20%, which allows for attributing it to qualitatively new flight controllability structures.

The paper analyzes the dynamics of a low-orbital space tether system (STS), consisting of a main and a small space vehicles, and a tether connecting them. Under consideration are the stages of deploying, free motion and stabilizing on a low and nearly circular orbit (170-180 km). The tether escapement is performed fr om the main spacecraft by the mechanism operating only on braking action, according to the feedback principle of measuring the tether length and escapement velocity. The tether length after deploying termination is several tens of kilometers.

The study of the STS motion on a low orbit becomes more complicated due to the necessity of considering the atmospheric drag acting on all system elements including the tether. It was demonstrated, that at the end of the STS deploying in a position close to vertical, unavoidable system oscillations relative to vertical occurred, caused by joint affecting of gravitational and aerodynamic forces (aerogradient effect).

The author suggests a nominal deploying program of the low-orbital STS at the position near to vertical. The proposed STS deploying program, compared to the known programs, accounts for the effect of the aerodynamic force acting on the end-bodies and the tether. The program law elaboration is realized by a simplified model with inextensible tether, and written in the orbital moving coordinate system. To verify the effectiveness of the suggested program the STS mathematical model with distributed parameters, wh ere the tether is represented as an aggregate of material points was elaborated and applied. Numerical simulation of the deploying process revealed that the suggested nominal program of the STS deployment allows decrease the amplitude of aerogradient oscillations of tether relative to the vertical by several times.

Simulation of the stages of free motion and stabilization was performed on the model with distributed parameters. When the orbital height of the system's center of mass decreases to a certain value, the low-orbital STS will switch to the stabilization motion in a given range of orbital height (170-180 km). Stabilization of the system orbital motion is realized by a correcting thruster, located on the main spacecraft. Employing the correctiing thruster ensures the flight stabilization of the low-orbital STS in the given range of orbital height. At the stage of STS motion stabilization, restrictions, imposed on the tethers angle deviations from the vertical are executed.

The safety of aircraft landing on approaching the landing strip in difficult weather conditions is associated not only with the need to create light and strong devices, but also, mainly, the search for new principles (methods and tools) for building control systems, since the aircraft landing is the most laborious process and largely unsolved problem to date.

Safety upgrading is achieved by control automating while approaching the landing strip and aircraft landing. It is obvious that the use of standard methods for modeling, analyzing and managing of complex multi-level systems becomes less possible with complexity increasing. In this situation, fuzzy control methods are the most applicable to such complex technological processes as the control of aircraft landing.

Aircraft control systems based on the principles of fuzzy logic, allow increase the course stability of the aircraft. In such situations, energy consumption is reduced and the response time of the system is increased simultaneously. Besides, it is possible to make the system as a whole more stable to the effect of disturbing factors compared to the traditional aircraft automatic control systems.

Practice shows that the operator, in conditions of good meteorological visibility range, satisfactorily lands an aircraft without the help of a program control system and a trajectory control system.

In the case of poor meteorological visibility, with the lack of visual contact with the runway strip, radio technical, optoelectronic and inertial navigation systems are employed for aircraft landing. They are used in the control system as sensors of primary information for the automatic control system (ACS). Such systems are termed course-glissade systems. They determine the position of an aircraft on the course and on the glide path.

But, even with modern control systems provision equipped with computerized hardware and software systems, which functionality is largely determined by software, applied diagnostic models, information processing algorithms, etc., the final decision-making is delegated to the human, which is a consequence of the insufficient effectiveness of diagnostic models, reflecting real ACS and the environment.

Thus, the structure schemes of similar systems in the following stages should include the links with fuzzy transfer function W_N(p) instead of links with functions W_o(p ) or W_a(p ) . To this effect, it is rational to implement the of the operator's behavior in such a situation as the basis for the fuzzy controller synthesis. In this situation, namely the methods of fuzzy control are the most applicable to such complex technological processes that will allow reduce by 10 times the duration of the longitudinal and horizontal movements' transients. The pilot in this case operates as a controller for the state of the control system.

Thus, the task consisted in developing models and algorithms for the design of control systems based on the methods of the theory of fuzzy-multiple apparatus.

A program in the C++ programming language was created to reproduce the min and max operations in on-board systems for automatic control of the aircraft lateral movement with applicaiton of fuzzy logic.

The article presents the developed method for efficiency increase of the operator, performing traffic surveillance by an unmanned aerial vehicle (UAV) with the built-in computer vision system. Analyzing video information, received via the radio channel in real time mode by the human-controlled point, is associated with errors in decision-making. These errors are stipulated by the vast volume of information, which overburdens the operator, and, as a rule, by the so-called human factor. Productivity of such system can be increased significantly through addition of autonomous road situation estimation system. The UAVs equipped with surveillance systems, such as video cameras, receive images onboard (video sequences), and are able to extract from them the objects of interest: roads and transport means.

Estimation and analysis in this article are ensured by the road incidents consequences severity classification. The work employs the classification consisted of five classes. Each situation class is described by attributes' dictionary, which separates the attribute space into non-crossing areas, corresponding to the selected classes.

In addition, the article describes the developed hierarchical structure of “Description of the Scene Being Surveyed”. This structure relates to the so-called semantic descriptions, is rather universal, and ensures the possibility to describe various road traffic situations.

The article presents the technique for traffic situations classification over the images. It demonstrates the example of the situation classification based on the real image of the road accident.

The presented article deals with the problem of semi-automatic initialization of the selected object tracking by unmanned averial vehicles (UAVs) or drones. Here, we proposed an algorithm of the position and sizes refining of the boundary rectangle of the tracked object at the start time (on the first frame) based on saliency detection algorithm, which simulates the map of human attention. The advantage of the proposed approach is that it applies the principles used by the human visual system: the color contrast, the main attention is centered on the central objects. The first stage of the proposed approach consists in preliminary image processing (noise removal) by the Gaussian filter and converting the image into the CIE LAB color space. The next stage is segmenting the image into homogeneous areas (superpixels) by simple linear iterative clustering (SLIC) algorithm. Undirected graph is employed as a container for information on segments storage. Based on information from the resulting graph, measures of identity, which assign superpixels to the background or an object, are computed. The resulting saliency measure is computed for each superpixel by optimizing the target cost function, which combines the measures of identity to the background, an the object and the smoothing component. The obtained saliency map of the image superpixels is binarized by the Otsu method. After that, the pixels belonging to the shadow can be determined. At the final stage, the operations of morphological filtering were applied to reduce fragmentation of objects and an algorithm for allocating coherent components, assigning the final dimensions and position of the object of interest for tracking initialization.

The algorithm was used to initialize a number of fast and effective methods of object tracking: DCF_CA, MOSSE_CA, SAMF, DCF, DSST, MOSSE, SRDCF.At the same time, the quality of the tracking was tested on the largest and most complex database of video clips, shot from an unmanned aerial vehicle – UAV 123.

The results of experimental testing allow conclude that the best tracking quality as a result of initialization by the proposed algorithm is achieved by tracking algorithms “SRDCF” and “MOSSE_CA”. In assessing the performance, you can notice that “MOSSE_CA” tracking algorithm is noticeably superior to the other algorithms. In this way, the most suitable algorithm for tracking objects by UAV, along with the proposed initialization algorithm, is “MOSSE_CA”, due to its least sensitivity to the of initial initialization accuracy and fastness among competitors.

The proposed algorithm does not require special hardware and can work in real-time. It is implemented in C ++. The average time required refining the object, occupying 40% of the image size of 256 × 256 pixels, equals 60 milliseconds on the Intel^® CoreTM i5-3470 CPU @ 3.20GHz.

Popular in recent decades concepts of an “all-electric aircraft” and “more-electric aircraft” assume full or partial replacement of centralized hydraulic systems by centralized electrical systems, and hence application electromechanical and electro-hydrostatic actuators alongside with electrohydraulic steering actuators. In the aircraft of the abovementioned class of the nearest perspective, the electric actuators with hydrostatic transmission are more applicable for the main steering surfaces controll. It is associated with higher indicators of wear resistance and non-failure operation of hydraulic cylinders. Electromechanical actuators application for these purposes is constrained by their insufficient reliability.

The article proposes solutions aimed at increasing the reliability of electromechanical actuators by both element-by-element and structural redundancy. An obligatory element of an electromechanical actuator optimized for weight and size indicators is a mechanical gearbox, which can wedge while operation. Multichannel electromechanical actuators can be constructed by one of the considered schemes, free of gearbox wedging. According to these schemes, each actuator channel must contain a motor shaft locking clutch, employed in case of a channel failure, and while channels serial operation as well. Alternative option are the schemes, requiring employing of clutches splintering the faulty channel off the common load – the steering surface. Such clutches should have reliability indices higher, than those required for an actuator all-in-all. The authors propose to construct them based on low power electromechanical actuators with redundancy.

Based on the comparative analysis results of the schemes options for constructing an electromechanical steering actuator with redundancy, three basic schemes were defined for which the preliminary failure rates were calculated.

The results of calculations allow us to consider the basic schemes of a electromechanical actuator with redundancy as an alternative to electro-hydrostatic steering actuators for a primary flight control system.

The purpose of the study of this scientific article is accuracy improving of monitoring robotic complexes positioning in surrounding space while performing various types of motion.

The subject of the study are methods for joint navigational information processing obtained by on cartographic and instrumental data.

The article analyzes the approach and problems of high-precision positioning of unmanned aerial vehicles and ground-based robotic complexes in the surrounding space.

The initial data of the study are based on formation of local areas of the Earth surface employing cartographic data and instrumental measurements. The article presents the main stages of the methodology for the map high-precision local areas formation by mathematical processing of redundant navigation parameters at the base points.

The methodological approach differs from the known ones by the presence of correlations between the map errors and allows the accuracy increase of navigation parameters determination over the entire area of the local section by 1 m, and by 3 to 5 m at base points.

The studies can find application in various fields:

— when solving problems of high-precision positioning of air and ground robotic complexes in the surrounding space;

— when solving the problems of ensuring and developing the Earth Deformation Control Service within the framework of the Federal System of Seismological Observations, particularly in the highlands of the country;

— rapid creation of a multi-information cartographic basis of various scales with account for correlation dependence of navigation and geophysical information.

Aerospace industry is in special want of high-efficiency electrical drives (motors), which allow reduce electric energy losses and rise productiveness of equipment. The number of electric drives onboard an aircraft varies from 50 to 220 pieces. With high tech development, the number of electric drives onboard an aircraft will only grow, and insignificant efficiency increase of all electric drives in the aggregate will lead to significant fuel savings. Three-phase induction motors with squirrel-cage rotor are in most common use in aerospace industry as fuel transfer drives. Asynchronous motors with maximum possible energy characteristics possess an efficiency below 80% and a power factor below 0.82. A possible alternative to asynchronous motors are BLCD motors, though their employing as pump drives becomes rather hindered due to cost intensive control system and large weight and size parameters. Another possible alternative to asynchronous motors may be a synchronous motor with permanent incorporated magnets and direct asynchronous start-up. The article is devoted to the analysis of structures of synchronous motors with incorporated permanent magnets and asynchronous start-up, for fuel-transfer pumps drives for the aerospace industry. The analysis was performed by computer simulation in the Ansoft Maxwell software package. The article proves the superiority of such motors over asynchronous motors. The structure of synchronous motor with incorporated permanent magnets and asynchronous start-up, which meet the requirements to fuel transfer drives for aerospace industry, was obtained based on computer simulation. The obtained results can be employed for the design of synchronous motors with permanent incorporated magnets and asynchronous start-up.

The promising concept of all-electric aircraft free from pneumo- and hydro actua-tors for flight control and stabilizing rotation frequency of main starter-generators supposes significant rise of power supply capacity up to 1.5 MW and more. To ensure high reliability indices and quality of supplied electric energy, the parallel con-nection of supply channels of onboard electric power system should be provided, as well as reversible (bidirectional) interconnection with stand-by low-voltage batteries.

To realize the concept of all-electric aircraft, the article suggests application of the so-called differential higher voltage DC-link with frame grounded averaged-potential (“zero”) wire.

Apart from the well-known benefits of the high DC voltage distribution subsystems, suggested high voltage DC-link has specific benefits, which allow substantiate particular requirements to the power supply systems for domestic all-electric aircraft.

As an example, the article presents the power circuit of the electric power supply combined channel with high voltage DC-link and standby battery based uninterruptible source for combined electric power systems with modular-scalable architecture. It also describes this channels operation. The reviewed structure of a single-phase power supply channel with high voltage DC-link may be recommended as an interrelated group of unified modules of switched mode converters applicable for synthesis of combined power supply systems with modular-scalable architecture and enhanced power supply capacity, and for all-electric aircraft power supply systems in particular.

The article suggest also the combined power supply system with distributed higher voltage DC-link and modular-scalable architecture for all-electric aircraft. Schematic and algorithmic solutions for three types of multifunctional switched mode converters, encompassing all specter of necessary conversions in onboard systems are considered. These solutions allow realize power supply systems with modular-scalable architecture for all-electric aircraft with account for import substitution of power electronics product range. The article presents the example of a simplified combined power supply system with differential higher voltage DC-link of hypothetical all-electric aircraft with four cruise engines and four main starter-generators.

The main issue of this article is modeling and reviewing of the possibilities of high-speed induction motors to modernize conventional and newly developed technical solutions.

The performed analysis reveals that by the year 2050 electrical energy production and consumption will practically double relative to 2015. Electric drive is one of its key consumers, and induction motors as the main motors. In this regard, presently the issue of meeting the rapidly developing industry requirements for developing highly effective reliable models capable of operating under conditions of drastically changing load arises.

The prospective arrangement to be modernized with such electric drive is a gearbox of a gas turbine engine. Difficultly controlled and tightly coupled with gearing-system, the pumps can be substituted by their lightweight, small-sized analogs in the form of electric drives. The authors offer two possible structures of modernization of various degree of complexity.

The induction motor modeling is complicated by computing parameters of its equivalent scheme. The article presents the review of the key values hard to calculate, and simplifications description, which were assumed while those parameters computing. The induction squirrel-cage motor with two rotor windings was selected as the basic model.

Two models for studying characteristics of high-speed motors were developed with MATLAB-Simulink. The first model simulates the motor with frequency regulator, and the second one is finished electrically driven gearbox of gas turbine engine aggregates. The problems of harmonic components generation by frequency converters are considered as well.

High frequency motors simulation results were compared to series-produced analog. They demonstrate the superiority of the new models compared to conventional, such as less jitter of the velocity curve, reduced inrush current, faster transients and increased torque. Comparison of variable-frequency control technique advantages with series-produced analogs was performed in the final part. The wider capabilities of the enhanced frequency range are demonstrated.

The energy path separation of switched mode electric energy converters into several channels with power switches control in different phases is a promising method for increasing the energy efficiency, reliability, and manufacturability of these devices. Selection of pulse-width modulator control mode for power converting cells while synthesizing the structure of a switched mode power converter with stabilized output voltage is of fundamental importance. Current mode has a number of significant advantages over Voltage mode.

They are as follows:

– better regulation dynamics;

– possibility of simpler overcurrent protection ensuring;

– automatic uniform distribution of currents between power converting cells. The main disadvantage of current mode is the possibility of subharmonic oscillations occurrence in continuous current mode. To ensure subharmonic stability, slope-compensation, or the duty cycle limited within the range of 0–0.5 are applied.

The article proposes three circuit solutions for multiphase basic frequency generator with a duty cycle equal to 0.5 and uniform time shift between phases for multiphase pulse-width modulators in Current mode.

A generator with a number of phases of N, defined as N = 2^k, where k is a natural number, can be built employing T-flip-flops. The N-phase generator circuit, requires N–1 triggers. The disadvantage of the scheme

is the limited choice of the possible number of phases. Its advantages consist in realization simplicity and automatic restoration after a failure caused by external jamming.

A generator with any integer number of phases N – k can be realized employing a shift register. N-phase generator circuit requires an N bits register. The circuit is also insensitive to failures.

A generator with an even number of phases N – 2k can also be implemented employing a shift register. To obtain N phases according to this principle l – N/2 register bits are sufficient. The drawback of the last proposed scheme is inability of its automatic recover after a failure.

The article presents the results of statistical studies of mechanical properties of the deformed semi-finished products from Ti-10V-2Fe-3Al titanium alloy based on analysis of literature, experimental and commercial data, by the “Stadia 7” software package. The effect of reheat temperature for quenching T_h, as well as ageing temperature T_ag on mechanical properties was evaluated by method of regressive analysis of the published tе data. Equations for computing polymorphic transformation temperature beta-transus temperature β_tr and the quantity of primary α-phase, formed while quenching process on the temperatures interval from 700 °C to β_tr were obtained:

β_tr = 890 + 22,3Al - 13,9V - 8,0Fe,

n_α = (0,3 ± 0,02)·( βtr-Т_h), %

.

1608 ingots and die forgings, manufactured by the industrial technology in 2007-2016 were also the subjects of research. All forgings were subjected to thermal treatment, consisted of quenching (763-798°C for three hours followed by water cooling) and ageing (500-515°C for 8 hours followed by air cooling).

The following factors were selected for statistical analysis: alloying elements’ and impurities’ content, beta-transus temperature, alloy structural equivalents in aluminum    and molybdenum, hardening temperature T_h and the aging temperature T_ag, the mechanical properties (offset yield strength σ_0,2, tensile strength σ_в, elongation δ , reduction of area ψ , fracture toughness K_1C). Primary statistical processing and correlation-regression analysis were performed.

Correlations between mechanical properties with deviations of the brand composition and heat treatment modes were established. At the first stage, pairwise correlations between the investigated factors were analyzed. The results of the analysis revealed that each element separately either does not affect, or affects weakly the level of mechanical properties of forgings, which is most likely stipulated by small intervals of their change. The joint action of the elements, which was evaluated by and , appeared to be more significant, the coefficient of multiple correlation was R=0.3-0.5, the fraction of the of the properties variation was γ ≈10-25%. Coefficients of multiple property correlation with quenching temperature and aging temperature were equal to R = 0,3-0,6 depending on the year of production. The joint effect on theproperties of all four factors (, ,Т_h, Т_ag) is evaluated by the coefficients R = 0,35-0,67, γ ≈ 12-45 %. The rest of the variation is stipulated by factors that could not be determined based on the data studied. The generalized regression dependence of the tensile strength of Ti-10-2-3 forgings on the chemical composition and heat treatment modes is:

The assumption on the possibility of employing interference pattern of aluminum alloys surface in polarized light for determining crystallographic orientation of separate grains was put forward. This assumption was tested on the example of aluminum alloy AD1. Optimum modes of electrolytic etching of AD1 alloy, under which the grains' boundaries were sharply defined, and necessary interference pattern of the grained structure was attained, were defined. Electrolytic etching was being performed in a 40% solution of hydrofluoric acid, boric acid and distilled water at 1.7-1.9 A, 100-110 V, and etching time duration of two minutes. It was established that the interference pattern of the sample surface changes with prolonged exposure in the open air. This was due to the oxide film's growth process. Employing literature data on elasticity modulus of aluminum mono crystals depending on crystallographic direction, the article defines the relationship between the grain elasticity modulus and its crystallographic orientation over three directions by the scanning probe microscopy method using NanoScan-3D device. Scanning of the studied section with a size of 128 х 128 µ m was carried out at a speed of 30 µ m/sec.During the scanning process, the signal from the indentation sensor was recorded and processed, resulting in a surface profile map (Z_opt). Modulus of elasticity of separate grains was determined by the method of removing the curves of the indenter's supply to the surface of the sample for each grain in the section under study.

While comparing the interference pattern with the distribution of modulus values, it was found that the grains of blue color corresponded to minimum values of modulus of elasticity from 46 to 55 GPa. Maximum values of modulus of elasticity were in the range from 69 to 78 GPa, and corresponded to yellow grains. It was established also that pale orange grains correspond to modulus of elasticity from 55 to 64 GPa. As a result, the assumption was made that the blue grains have a crystallographic direction [100], since they have minimum modulus of elasticity in the array of obtained values. Yellow grains have a crystallographic orientation [111] and maximum modulus of elasticity from the values obtained. Pale orange grains occupied an intermediate position by value of modulus of elasticity, so it was assumed that their crystallographic direction corresponds to [110].

The developed technique is characterized by simplicity, low energy intensity, and less time consuming, in contrast to the methods traditionally used for this purpose. This technique can also be employed to determine the crystallographic orientation of individual grains of other aluminum alloys.

The aviation industry is one of the most high-tech industries not only in the product design and development, but also in the production process posing high requirements on personnel qualifications. The system of qualifications assessment and certification in the aviation industry helps to solve the issue of staffing of the production process due to:

• Reconcilement of employers' requirements to graduates' qualification;

• Independent and objective assessment of the qualifications' mastering level;

• Accreditation of educational programs by employers.

The systems of professional certification in Europe and the United States has been functioning since the 1980s. In 2007 a large-scale activities on creation of National system of qualification started in Russia under the auspices of the RSPP. Since 2014, this work was being performed on the ground of the National Council on Professional Qualification under the President of the Russian Federation. The Council for professional qualifications in the field of aviation was established in 2016, it included representatives of major employers and their associations, public authorities and educational institutions.

At the first meeting of the Council, it was decided to consider the possibilities of employing and adaptation of the project sectoral qualifications frameworks in mechanical engineering for aircraft industry. For this purpose there are all preconditions, since the enterprises need the skilled workers in the first place, and it was for them that the Sectorial Council on Machine Building develops qualification requirements.

Based on Federal law No. 238-FZ “On independent qualification assessment”, 108 organizations, getting the status of centers of assessment of qualifications were selected, in which more than five thousand people have already confirmed their skills.

Currently CTCS in the field of aviation has not been not established, however the need for its creating is urgent. Though MAI cannot act as the CSC organizer, it is likely expedient to enlist the services of the MAI teaching staff to its activities In particular, the administrative tasks of CTCS can be transferred to MAI, which has extensive experience of performing similar procedures: admission company, examinations, etc.

The proposed system will help the industry economically, as well as strengthen the ties between the labor market and educational sphere.

The problem of Russian air companies' transition from employing foreign-made passenger aircraft to domestic ones is under consideration. The article analyzes the status of passenger aircraft being under operation or being ordered for the future. It also defines the aircraft ownership i.e. financial leasing, or operational leasing or airline's property. Мost attention is payed to the study of an aircraft operational leasing, since regulation of aircraft park being in temporary service will allow release market niches for domestic built aircraft.

The article presents a methodical approach, allowing assess variants of freight capacities volume formation based on statistics monitoring to implement new airplanes of domestic manufacturing.

Analysis of the rules of statistical accounting applied by US airlines led to the idea of the above saic proposed methodology. The Bureau of Transportation Statistics (BTS) of the USA publishes monthly the detailed airlines reports on distribution of aircraft types by airlines, reflecting distance factors, the number of planned and actually performed flights, passenger traffic, cargo and mail, available seat-miles, passenger-miles, ton-miles, flight hours and aviation fuel consumption. Based on the detailed data, brief reports on aircraft employing are being compiled and can be sent to aviation organizations, such as ICAO. The detailed presentation of information allows perform studies on modernization of aircraft fleet under operation adequately and without extra resources.

The air transport of Russian Federation publishes brief statistical forms on aircraft availability and usage (32 civil aviation and 33 civil aviation). To obtain detailed data on performed air service by airplanes of airlines the data on full schedules (SRS Analyser) and passenger transportation along the routes are being integrated.

The calculations simulating the workout resources of an aircraft in use are performed using the network modelled in such a way. In the longer run, the demand for airlifts rises, the aircraft in service drops out, and a niche of free seat-volumes for new aircraft implementation appears.

The source of information is the Transport Clearing House statistical database on aircraft fleet at disposal and employing this fleet by airlines and industry at large, transportation between pairs of cities, as well as international databases. The Flight Global database is used to analyze the state of passenger airplanes' park. It gives a comprehensive idea of air transporters' airplanes under operation and aircraft building industry perspectives. The airplane schedule and freight capacity were accepted according to SRS Analyser database. In calculation for perspective, the United Aircraft Corporation plans on aircraft building up to 2037 were accounted for. The market niche formation of new aircraft implementation is affected mainly by the demand on passenger aircraft seats as a whole and by segments; terms of passenger airplane disposal; signed contracts of airlines on the delivery of foreign-made aircraft; delivery plans of domestic manufactured aircraft.

Presently it is necessary to develop highly effective assessment of high-tech enterprises economic security level as each aircraft building enterprise needs integrated self-concept of its production-commercial and financial-economic activities.

The main objective of assessment of the level of production activity economic security is assessment of risks on integrated system of indicators accounting for specific branch features at the enterprises of high-tech industries. The authors recommend employ the hi-tech enterprises' risks assessment based on a qualitative or interval method. Internal self-concept of state of production of aeronautical engineering development, and assessment of production development dynamics of the enterprise should be the main task of high-tech production monitoring.

The internal self-assessment of a production condition of creation of the aircraft equipment and assessment of dynamics of development of production of the enterprise have to be the main objective of monitoring of hi-tech enterprise. For this purpose, it is necessary to employ the technique, which would describe the main approaches and basic self-concept procedures for the risks of production departments of high-tech enterprises.

The article considered and developed managing system for high-tech production with account for economic security requirements, which will allow ensuring the raise of high-tech products competitiveness.It suggests assessment indicators for probability category of economic security factors coming-in, risks register, and risks map in high-tech branches of the industry.

Developing the system of economic security of production activity in aircraft building will allow forecast the aftermath of internal and external hazards on both production processes, and high-tech enterprises' activities at large. Implementation of the above said procedures at the aircraft building enterprises will allow ensure the necessary level of economic security, as well as optimize production processes at the enterprises of high-tech industry branches.

Enterprises are constantly facing competition, both at domestic and foreign market. The competition promotes the development of export, but the risks associated with an unsatisfactory estimation of activity of an enterprise and improper planning may inflict damage to the company.

To consider the issue of products export one can employ the following method of export potential assessment to promote products and services at foreign markets:

1. Assess the popularity of manufactured products or services at the domestic market. If they are successfully sold in the local market, they will be probably in demand abroad, at least at the markets of the countries with similar socio-economic conditions and needs;

2. Evaluate the unique or most important features of the produced goods and services. If they are hard to be reproduced abroad, there is a possibility the company will enjoy the success, as unique goods do not face the severe competition and the demand for them is high.

Since the Holding Company “Helicopters of Russia” occupies the leading positions in military-industrial complex development, the article is devoted to the study of export potential of one of its enterprises, namely, Arseniev aircraft company “Progress”, which, plans to export military equipment in 2017.

To avoid the above said risks, the company was proposed to employ the model developed by the author for export potential predicting, based of complex evaluation of enterprises. The model is based on evaluation of the basic technical and economic, accounting and financial indicators, as well as indicators of the enterprise's management and its scientific and technical potential. It includes also statistics of qualitative and quantitative structure of personnel, gender and age structure, work experience of engineering personnel, managers, fellow laborers, workers, as well as information on the research work and capabilities of research projects performing.

Summarizing the results on the above listed factors, a model of the multiple regression describing the dependence of the export potential of human resources was built, which was subsequently automated with a software application.

As the result, the program was:

– effective, since the payback period was 45 days at a low annual economic effect;

– universal, since applying this tool is possible for other enterprises (not only the aircraft industry);

– easy-to-use, since the users need only necessary information on the enterprise dynamics, and, according to the theory of mathematical modeling and regression analysis, the more data will be used, the higher the model adequacy and forecast accuracy will be.

Cylindrical shells are the most common structural elements of rocket engines. When loading by the temperature gradient in radial direction radial temperature stresses occur in them. Such stresses in carbon-carbon shells can be rather dangerous notwithstanding that they are much smaller than the circumferential and axial ones. Moreover, they substantially depend on the thermal conductivity of the carbon fiber material and the shell structure.

The article suggests the equation for the structural thermal conductivity (contact thermal exchange) evaluation of a cylindrical shell in radial direction. When calculating with the equation the carbon fibers' roughness was not accounted for due to the presence of pre-preg matrix, and the shell was divided conditionally through-the-thickness into several layers. The contact forces acting on the fibers were determined based on a primary evaluation of the temperature stresses. The results of the shells' made of carbon fibers calculations with a diameter of 0.02, 0.05, 0.2, 0.5, and 1 mm are presented in the form of tables and graphical dependencies. It is shown, that the elasticity modulus of the first genus of carbon fibers' surface layers can be accounted for in the calculations. It was revealed, that in shells with internal warming-up the specific pressures at the areas of contact spots of the adjoining fibers could reach several tens of kilograms per square millimeter. There is a risk of the carbon fibers structure stratification in the shells with the external warming-up. Thus, we recommend conduct tensile or bending tests with small-sized samples, cut from the shell in radial direction. Tests of such samples should be carried out according to the methodological instructions.

One of the ways to increase the aerodynamic quality of modern and prospective aircraft consists in wing aspect ratio increasing. Such increasing leads to the occurrence of various new aspects of the structure loading, strength and aeroelasticity. One of these aspects is increasing of the wing flexibility, and as a consequence, possible in-flight structures deformations effect on aeroelasticity characteristics.

The paper presents a review of publications on the deformation effect on various aeroelasticity characteristics. It suggests and substantiates a computational method for studying the effect of quasi-static deformations of the wing on static and dynamic aeroelasticity of an aircraft. This method is based on automated generation of a set of aircraft computational models using the Ritz polynomial method. The paper presents the examples of a wing in-flight deformations effect on characteristics of static elasticity, frequency and shape of elastic vibrations, and flutter characteristics.

The results of the developed method application for aircraft of various configuration allowed establishing the main regularities of the effect of structure's deformation on aeroelasticity characteristics.

The effect of in-flight deformations on the characteristics of static aeroelasticity and load is determined by: 1) effective wing span decrease; 2) aerodynamic forces direction changing; 3) increase of the effective dihedral angle. Characteristics of longitudinal motion can reduce by 5-6%, while characteristics of lateral motion can increase or decrease by 5-15%.

The dynamic aeroelasticity characteristics change is determined mainly by the increase in the interaction of torsional oscillations of a wing with bending vibrations in the chord plane. For the unmanned aerial vehicles with a wing of extremely high aspect ratio, this effect can lead to a significant decrease in flutter speed (up to 30-50%). For modern airliners, the decrease in flutter speed due to the in-flight deformation does not exceed a few percent and lies within the accuracy of numerical methods.

An important feature of the method is its integration into a multidisciplinary design complex ARGON, validated while solving aeroelasticity problems in many practical applications.

A comprehensive material on operation of spacecraft with solar electric propulsion systems is accumulated by now. The latter are designed for spacecraft correction on both geostationary and circular low earth orbits.

At the same time, there is a tendency to developing of small spacecraft of various purposes, such as scientific, communication, Earth remote probing, navigation, hydro-meteorological etc. operating on low circular orbits with the height within the rage of 180–280 km. Such spacecraft are relatively cheap and possess the mass of 10 to 500 kg. However, such indicator as minimum orbit height, its relationship with the spacecraft weight and size, as well as parameters of its engine unit remain undetermined. System analysis and experimental data on spacecraft with solar electric propulsion systems, operating at the height of 140–280 km are practically inaccessible.

The paper considers the problem of small spacecraft transition fr om a higher circular orbit to a lower one. As far as the Earth atmosphere gradually transfers to vacuum, the aerodynamic drag force grows while a spacecraft descent. We suggest surmounting this force through the electric jet engine thrust power. It is obvious that while the spacecraft descent the aerodynamic drag grows, and such parameters as thrust power and electric jet engine power should be increased continuously. At large the problem becomes dynamical. Besides, the main cause of the orbit height limiting will be the drag force of the solar battery. Thus, the minimum orbit height h_min below which the spacecraft, equipped with the solar electric jet engine cannot exist, is limited by the spacecraft drag force due to the solar battery. At the lower altitude the battery's drag force will be greater than the electric rocket engine thrust force.

For the spacecraft motion analysis, we assumed that the solar battery takes the shape of an autonomous panel with rotation angle control to the sun radiation direction. The power flux density or the solar radiation at the Earth orbit is Q = 1400 W/m² (solar constant). The efficiency of photoelectric transducers based on a three-stage gallium arsenide (GaAs) equals to η_SB= 0.22. The solar battery specific power is α = 308 W/m². If the solar battery plane is oriented normally to the orbital movement velocity vector the drag factor equals to C_SB = 2.15, and if it is oriented along this vector it equals C_SB = 0.15.

If an ion thruster is used as an electric jet engine its specific impulse is assumed as Ι_SP = 4500 s, and its efficiency equals to η_T= 0.7. In case of plasma engine of the SPT type I_SP = 1700 s and η_T= 0.55 correspondingly.

The lower limits of the orbit altitude h_min = 200 for the solar electric jet system with the ion engine, and h_min = 180 km with plasma engine of the SPT type were established by the results of the performed analysis. The upper lim it of the altitudes descending from which requires continuous build-up of the electric jet system solar battery area to overcome the atmospheric aerodynamic drag is the altitude of h_max = 260 km.

The paper demonstrates that for a spacecraft continuous exploitation at the latitude of 180–260 km application of solar electric jet engine and atmospheric gas as a working agent is possible. Application of high- frequency ion engine of 4.4–5 kW is expedient for the propulsion installation such kind. With the specified power and solar battery weight, the weight of electric jet propulsion installation will be no less than 90–100 kg, and the total minimum weight of the spacecraft will be no less than 500–600 kg.

With turbojet engine thrust reduction, its small size begins affecting the effectiveness on its elements. Lower airflow rate results in blades size decrease and relative radial clearance increase. It affects the efficiency of axial turbo-machines. Due to this, radial and centrifugal turbo-machines become more effective at small thrust values. The main goal of this study consists in determining the most effective structural scheme of a turbojet engine for the thrust range from 0.1 kN to 100 kN. The problem was solved by performing the engine multi-criteria optimization employing ASTRA CAD, developed in Samara National Research University. The total weight of a power plant and fuel, as well as specific fuel consumption were selected as performance criteria. The optimized variables are the gas temperature prior to the turbine, and total pressure ratio. According to the optimization results the following inferences were drawn. With optimization of the engines with the thrust, lower than 25 kN, corrections on their small-size should be accounted for. With the engine thrust decrease, the optimal parameters of the working process are decreasing either, and the regions of compromises are contracting. The axial compressor is optimal for the thrust of 7 kN and higher, and with thrust decrease up to 1.3 kN, compressor of axial-centrifugal type becomes more appropriate. The axial turbine is effective up to 0.7 kN thrust value, and radial turbine is effective for small engines with lower thrust.

The article deals with the flight thrust determining method of a bypass engine with flows mixing in the presence of a non-uniform total pressure field at its inlet. The non-uniformity impact is taken into account for both air consumption due conventionally averaged total pressure at the inlet, and the specific thrust due to the overall pressure level reduction along the engine passage, and, respectively, the available differential pressure in the jet nozzle.

Earlier, the authors developed and patented the engine thrust determining method allowing evaluate its thrust while in flight under condition of the uniform flow at its inlet according to the measured operating conditions and external environment parameters. The presented work extends this simplified engineering method to the real case of a non-uniform total pressure field at the engine inlet. Moreover, it employs corrected values of the total pressure along the engine passage to compute the thrust.

Thus obtained, the value of the flight thrust can be used in both automated control system for its possible in-flight correction, such as partial or full flight thrust value restoration, and the complex engine diagnostics system to evaluate its deterioration rate and deterioration in performance of its separate parts and elements.

Calculated evaluations performed according the developed method with account for typical input total pressure non-uniformity revealed that the expected thrust fall will be of 8.9%, with about 8% herewith due to the air consumption reduction, and the rest is due to specific thrust decrease.

Recently, more and more attention is paid to the problem of ensuring unmanned flying vehicles (UAV's) invisibility in various frequency ranges due to the wide application of the systems with small sized UAVs for solving special assignment tasks. To ensure the UAV's invisibility in the audible frequency range at the specified distance from the observer in conditions of known terrain of application, the qualitative data on the UAV acoustic characteristics is required.

The experimental study of the small sized UAV's “Ptero-G0” acoustic characteristics was performed within the framework of the presented work. The UAV's power plant consisted of a single-cylinder gasoline internal combustion engine (ICE) and a small sized two-blade propeller with the fixed pitch. The acoustic tests were performed in TsAGI unechoic chamber AK-2.

The following main results were obtained as a result of experimental researches.

1. Energy, spectral and spatial characteristics of acoustic fields of a small sized propeller and single-cylinder four-cycle gasoline engine were obtained.

2. The small sized propellers diameter effect on UAV's noise and signature characteristics was studied. Recommendations on acoustic signature reduction of the UAV “Ptero-G0” were elaborated. These recommendations were implemented and accounted for by the “AFM-Servers” company while developing new flying vehicles.

3. It was demonstrated that a cowl mounting on the engine without both vibration and acoustic insulation could lead to significant noise increase of the power plant.

4. The possibility of employing the empirical model while solving the problem of a single-cylinder four-stroke gasoline engine's noise evaluation was demonstrated.

Within the framework of the subject's of UAV acoustics development, the authors are planning to proceed this work in the following main trends.

1. Studying the effect of the power plant's noise shielding by the airframe elements on UAVs noise and signature characteristics.

2. Studying the noise caused by the UAV airframe flow-around.

3. Development of semi-empirical model of the small-sized propeller's noise.

4. Implementation of the existing computation methods for audibility and signature boundaries into practice of noise indices evaluation of various UAV's types.

5. Software development for UAV flight trajectories' plotting under known weather and landscape conditions without the ability to detect it both by ear and with acoustic location finder.

Thermal flows emitted by power sypply unit (PSU) components lead to their heating, which, in its turn, may lead to changes of their operating characteristics up to their failure. Thus, the temperatures of these components should be maintained within the ranges ensuring maintenance of their operating characteristics. For this purpose a preliminary simulation of thermal processes in the PSU housing was performed.

The article presents a model for temperature distribution calculation in separate components of a radio frequency ion thruster's structure. These calculations were performed using ANSYS bundled software.

Due to the negligible effect of thermal flows from the thruster unit on the thermal state of PSU, thermal simulations of the thruster unit and the PSU were performed separately. The aluminum thermostatically controlled mounting flange, located above the gas-discharge chamber presents the boundary.

All PSU's structural elements in the computer model are simulated as simple geometric forms, such as cylinders or parallelepipeds with appropriate geometrical dimensions.

The total heat emission in the PSU unit from all its constituting elements is taken equal to 66.4 W. This value corresponds to the operating mode of a low-power radio frequency ion thruster.

To intensify maximally the heat removal by radiation, the emissivity factor of 0.9 was attributed to all external surfaces of the PSU unit components.

To maximize radiant heat removal, the outer surfaces of elements of PSU were modeled with the emissivity of 0.9. To increase the conductive heat exchange, a partial PSU components' potting (gersil) was performed.

The calculation used the real thermal contact between adjacent surfaces with corresponding values of thermal junction resistance. A series of calculations was conducted for various the compound's thermal resistance values from 1.2 to 2.7 W/(m·K).

The figure below shows the dependence of the temperature of the most heated component of the structure under various thermal conductivity coefficients of the compound.

The requirements for the thermal conductivity of the compound for filling the PSU's PCBs were determined.

When using materials with thermal conductivity exceeding 1.7 W/(m·K), it is possible to ensure the permissible temperature of electronic components at a temperature of the mounting flange reaching 50°C.

The developed physico-mathematical model can be employed at the stage of the ion thruster preliminary designing.

Development of the SPT-50 thermal model, thermal calculations and study of the model sensitivity to changes and to various combinations of internal and external heat exchange parameters was carried out with account for the requirements of the Thermica software applications package (SAP) based on employing of isothermal elements method. The thermal model under development consists of 130 elements. The radiation couplings for the SPT-50 anode unit's thermal model were computed employing Thermica V4 SAP. To obtain the information on the thruster thermal state during thermal vacuum tests (TVT) it was equipped with temperature detectors, installed on the thruster in places with enough access to the surfaces for contact welding, glue and other ways of mounting. The SPT's thermal balance thermal vacuum and thermal cycling tests were performed. The thermal model correction with the testing results was realized by thermal calculations employing the developed thermal model. The calculations did not account for convective heat exchange (imitation of vacuum). The ambient temperature was set the same as the during testing, and SPT's optical and heat emission properties were set according to the operating mode during TVT.

The developed thruster thermal model, updated by testing results tests, allows analyze thermal processes inside the thruster in the places where installation of thermocouples is impossible. After the SPT-50 thermal model correction one can define the critical design elements, thermally affected by the thruster. Based on the thermal calculation results, the element of wire with critical temperature level has been defined, and this value approached maximum temperature value of 220°C. To decrease the wire temperature, we increase the wire core section area to enhance the heat sink from the wires critical element. The calculations revealed that the temperature of the SPT's critical elements does not exceed maximum admissible working temperature. It confirms correctness of the approaches to selection of thermal design and parameters of the thermal regulation system of the SPT-50 anode unit. The presented thermal model of the SPT-50 anode unit can be employed for developing other options of thermal and mounting interfaces for other discharge and magnetic thruster operating parameters.

The article presents the results of transients' in discharge circuit studies during SPT-140D plasma thruster starting while its operation together with power processing unit (PPU). SPT-140D is an electric thruster developed by the Design Bureau “Fakel”. This thruster was running on Xenon with the discharge voltage of 300 V and power of 4.5 kW, ensuring reactive thrust 280-290 mN and 1750 s specific impulse of thrust. At present, this thruster is ready for flight application for spacecraft motion control. The PPU unit was developed and manufactured by the Scientific and Production Center “Polus”. Since the main discharge is one of the powerful PPU loads, the main attention was payed to the study of transients in the power supply circuit of the main discharge. The obtained data was used for the development of imitation model of the named transients and electric imitator of the thruster for off-line PPU optimization and testing without the thruster. In addition, the information on the specifics of the thruster operation was obtained. The most interesting among them are the following:

1. In the course of the thruster starting, after the main discharge ignition by the discharge voltage increasing with the rate of about 1 V/ms, the main discharge could ignite by various discharge voltages. Though after the discharge ignition its parameters during various start-ups vary according to one and the same averaged dynamic volt-ampere characteristic, close to the “static” characteristic obtained with slow voltage changing.

2. Various oscillation modes of the discharge parameters were revealed, arousing at the various stages of discharge voltage variation, and changing drastically with the small variation of the discharge voltage. It allows evaluate the increment of their build-up.

3. After reaching the nominal discharge parameters, the dominating discharge current oscillation mode frequency is 15-20 kHz. After an hour of continuous operation it reaches the value of 27-27 kHz, and its further variation is insignificant. It can be explained by the discharge chamber heating resulting in Xenon atoms velocity increase, decrease of their drift time through the ionization layer and acceleration leading to the frequency increase according to ionization-drifting oscillations excitation model.

Thus, the employed methodology of the study is useful also for conducting physical research of the processes in the thruster.

The subjects for study are dampers of various masses installed under the platforms of turbine blades.

The research issue is prevention of turbine blades failures caused by higher level of variable stresses.

The goal of the work consists in experimental and computational definition of effectiveness of shock-absorbing insertions' masses (3.3 g, 4.7 g and 5.8 g) for variable stresses reduction in the full-size turbine wheel.

The methodology of the work includes two trends: computation and experimental. The computation trend is based on modeling the damper using MSC.Nastran contact elements and estimating the reduction of vibratory stresses, by integrating the equation of motion in the time-domain employing the standard non-linear integration procedure by the Newmark method. The effect of the insertion on vibration frequencies of the blade was also studied. The experimental trend is based on a comparative analysis of the amplitudes of vibratory stresses in the blades both with installed damper and without them. Tests are performed on the CIAM bench test (manufacturer is Test Devices company). The turbine wheel is assembled for testing in a special way: one sector of the wheel is damper free, and the rest three sectors were equipped with dampers of various masses. The blades were prepared with strain gages, and in each sector the blades with maximum response to external excitation from the air supplied to the test chamber were selected. Tests were carried out for an unheated wheel.

The calculations revealed that the most effective reduction of vibratory stresses in the blade occurs when the holddown pressure of the damper to the bottom surface of the blade platform are 200–800 N. Such forces for damper mass of 3.3 g were caused by centrifugal forces at rotational speeds of the wheel in the range 35–70 % of the maximum rotational speed; this range is 29–58% for the damper of 4.7 g, and for the damper of 5.8 g, it is 26–51%. The affect of dampers weighing 4.7 g and 5.8 g is ineffective, starting, respectively, from 90% and 82% of the maximum rotational speed. According to calculations, the damper with mass of 3.3 g allows reduce the vibratory stresses by 22% at a resonant mode at the 87% of maximal rotational speed.

The tests revealed that, in comparison with damper and without damper, the blade frequency with shock absorber of 3.3 g increased by also 16%, and the oscillations' amplitude decreased by 25%. This correlates satisfactorily with the computation data.

Conclusions were drawn that the calculated and experimental results in these studies showed, in general, a satisfactory agreement with respect to both the reduction of vibratory stresses and the change in the resonant frequency when a damper was installed. Some discrepancy between the calculated and experimental data on the effect on the vibratory stresses of “heavy” inserts of 4.7 g and 5.8 g may occur due to the assumptions in calculations, as well as to the errors in the experiment and processing of the test results. To evaluate the effect of the amplitudes of vibratory stresses in the blade without damper, stiffness and mass of the damper, as well as friction coefficients on the effectiveness of the damper to reduce the vibratory stresses in the blade, additional experimental and calculated studies are required.

The article presents the results of the study in elaboration of scientific discovery No 314 “Phenomenon of the abnormal high growth of thrust in the ejector process with pulsating active stream” performed in MAI. The possibility of pulsating jet impulse increase through ejectorless addition of gas mass both from the external atmospheric environment and used up gas was shown experimentally. It increases the meaningfulness of the discovery. The physics of the process of the used up gas mass in pulsating stream is based the well-known phenomenon of wave interaction of cyclic masses with various velocities of front and tail parts.

Calculating and experimental studies substantiated the capability of creating a nozzle with the spherical resonator-thrust amplifier for air-breathing jet engine with stationary fuel combustion. The mechanism of gas masses adding in oscillating process is shown. The thrust amplification at certain gas-dynamic and geometrical relationships herewith can make 1.5 and more.

Carrying out of experimental studies on a vacuum bench (imitation of space conditions) has confirmed effective exhaust gas mass addition that opens new capabilities for increasing the thrust efficiency of space jet engines. According the test results, the constructive recommendations on the improvement of working process are given.

The results of computation and design working out of implementation of the obtained effects in the nozzle with the resonator of an optimum configuration for conventional air-breathing jet engine without its mass and dimensions characteristics derating present a great practical interest. Conditions and recommendations on calculation are given.

Possible perspective trends of further studies on implementation of the obtained effects of thrust increase at the expense of exhaust gas mass addition in liquid-propellant rocket engine and solid-propellant rocket engine with spin detonation fuel combustion, as well as in a gas turbine engine are determined.

In the last decade, much attention has been paid to the studies conducted in the interests of mathematical modeling methods developing in 3D setup. It requires a detailed study of various computational meshes constructing methods and their effect on the obtained results.

The problem of the aerodynamic characteristics computation of the of a perspective blade with deflectable trailing edge profile is important for both the development of wind turbine blades, compressor design for advanced gas turbine engines, and aircraft structures.

The effect of the computational mesh is studied while mathematical modeling of the inflow of a subsonic flow onto the profiles of a perspective blade with a deviating trailing edge. Verification, the convergence and correctness checkup of the solutions obtained, as well as verification on tasks having reliable and detailed enough solutions are necessary.

The objectives of this article are as follows: determining the accuracy of the numerical solution of the aerodynamic profile of the perspective blade with the deflected trailing edge, and testing the computational mesh with the potential to achieve industrial applicability. The feature in common is the use of wall-adjacent blocks adapted to geometry, applying herewith various approaches for their coupling with the external mesh. Analysis of the solvers application employing the Cartesian mesh reveals also the necessity of constructing mesh layers adapted to the surface of the body.

Analysis of existing designs allows us to draw the following conclusions. A simple deflectable trailing edge increases the lifting force by increasing the curvature of the profile. This increases the pressure on the lower surface of the profile, as well as increases its load-bearing properties.

A mathematical model of the aerodynamic processes occurring on the profile surface of a perspective blade from the back deflected edge while its on flowing by a subsonic flow is suggested.

An acceptable correlation of the results of the calculations made using structured and hybrid meshes circuits was obtained. Analysis of the results of numerical simulation employing various meshes revealed that application the meshes under consideration considered allows obtain close results. The structured meshes applied herewith consume less computation time. Hence, we will use the structured meshes as the best way to solve the problem.

Thus, the proposed mathematical model and the first method of developing the mesh can be applied to determine the numerical solution accuracy of the problem of flow past the aerodynamic profile of a perspective blade or wing with a deflectable trailing edge, as well as the mesh testing

The problem of detection and control of rockets' launching and flying is topical from the viewpoint of functioning safety of various ground and aerial objects of both military and general assignment. At present, significant attention is paid to identification and tracking optimization of low-speed point objects' of various purposes. A method for spectrozonal detection and control of low-altitude rockets through the exhaust plume of a solid jet engine was developed and theoretically confirmed. The authors formulated new spectrozonal features for detection of launched low-altitude rockets, based on the well-known experimental results related to the study of spectral emission of rocket engines plume. To detect and control the low-altitude rockets a new spectrozonal feature, possessing experimental property useful for applying for on both axial and radial directions was formulated. The issues of identification optimization of low-speed low-altitude point objects under variable atmospheric conditions were also considered. The general mathematical problem of optimization of the entire cycle of optimization was formulated and solved. Its gist consists in achieving the maximum possible value of the averaged signal received from the object by the infrared identifier, through the accepted model of variation of atmosphere optical thickness. The article demonstrates that as applied to subsonic flying objects, such as cruise missiles, it is necessary to ensure direct proportion between atmosphere optical thickness and a certain time index. It should be considered herein that the minute scale of atmosphere optical density can be easily controlled, and it presents the result of rapid weather conditions changes due to natural or anthropogenic factors. A certain increase of the value of the functional, obtained by the performed optimization, can be interpreted as a possibility of a certain shift of the total time interval to the left. On this basis, the solution of the formulated optimization problem points to the possibility of realization of the much safe mode of an object detection and identification under revealed.

Some increase of accepted target functional caused by carried out optimization can be interpreted as possibility for some shift of whole time interval to left. Solution of formulated optimization task indicates the possibility of for safe regime for detection and identification of object upon revealed advantageous atmospheric conditions.

Onboard radars operating in pulse Doppler mode possess characteristic feature in the detection zone. This feature lies in the fact that at each point of the detection zone the aircraft has the sector of directions. Moving along these directions it, cannot be detected by the onboard Doppler radar. This sector is named as the sector of an aircraft concealed motion directions. Due to these features there are concealed trajectories, moving along which the aircraft becomes undetectable by the onboard Doppler radar, such as an aircraft airborne early warning radar (AEWR). Most of these concealed trajectories are curvilinear with variable curvature. The article is devoted to the study of the aircraft concealed movement trajectories curvature in the onboard Doppler radar detection zone. The study of the aircraft concealed movement trajectory curvature in the AEWR detection zone was carried out to evaluate the possibilities of flying over such trajectories with account for and aircraft maneuvering characteristics. The results of the study led to obtaining the equation for calculation the curvature of any concealed trajectory in any of its points. The equation allowing determine the shape and size of a region in which the movement over the concealed trajectory is impossible due to the fact that normal overload exceeds maximum aircraft operating overload. It was established that for the valid parameters of an aircraft movement and AEWR aircraft this region is located within the circle with radius not exceeding 10 km with its center coinciding the location of the AEWR aircraft. The region, where aircraft high manoeuvrability is required, presents utterly small portion of the detection zone of AEWR aircraft. Thus, the aircraft concealed movement is possible over concealed trajectory practically in the entire detection zone.

One of the main problems of automatic control theory can be formulated as optimal functional links' forming between information and energy. The basic principle while control systems design consists in designing such systems, which are able to transmit or convert information with specified timing and phasing-in characteristics under condition of power consumption minimization for the given control law realization.

The main power consumption relates to actuating mechanisms while synchronous transmission fr om control system to the control object with concurrent increase of the energy level. The energy level limitations in actuating mechanism affect significantly such dynamic characteristics as stability, accuracy and noise reduction.

The problem of parametrical synthesis of the rudder servo drive actuating mechanism for the UAV's aerodynamic control system is interpreted as a system design optimization problem. The quality criterion of parametrical optimization problem is maximum effective power delivered to the control object from the power source, necessary to fulfill the required, most tough from the power consumption view point, motion laws of the control object (aerodynamic rudder) under specified parameters of aerodynamic load. Graph-analytic solution of the problem is based on plotting the dependencies N_max(F), wh ere N_max is the maximum effective power value; F = M_max/W_max is the robustness value of the actuating mechanism mechanical characteristic; M_max and W_max are the maximum torque and maximum speed of the actuating mechanism.

These dependencies allow define the optimal parameters of the actuating mechanism ensuring the fulfillment of all control object's required laws of motion, provided the minimized energy consumption for their realization.

The industry demand for high-speed electric motors with rotation frequency of 48,000 rpm to 120,000 rpm and power of 5 to 250 kW increases from year to year. A number of technological problems exists herewith, which retards the high-speed electric motors market growth. These problems relate to the issues of their output voltage stabilization, which are solved by employing static converter, stator back magnetizing, or rectifier; ensuring bearing assembly reliability, which are solved by employing non-contact bearing assembly, as well as the problems of rotor heat evolution reduction. The latter are stipulated by the complexity of fast-rotating rotor's cooling.

To solve this problem, the article studies losses caused by eddy currents in permanent magnets of high-speed electromechanical energy converters.

The eddy currents losses in permanent magnets and rotor retaining shell are generated by spatial harmonics caused by electric motor structural specifics, stator serration, windings diagram and distribution ratio, as well as temporal harmonics, stipulated by the external circuit, such as inverter. Moreover,with the improper selection of the electric motor parameters eddy current losses may lead to the permanent magnets overheating and their demagnetizing under the effect of this overheating.

It is generally assumed, that the losses stipulated by temporal harmonics are higher than the losses caused by the spatial harmonics. This statement is valid only for a number of structural schemes of high-speed electric motors. For example, the electric motors with toothed windings feature significant spatial harmonics. And losses caused by these harmonics are higher than the losses caused by temporal harmonics.

It is found that with rotation speed increasing the losses in permanent magnets have maximum point, after which they start decreasing. This is explained by the fact that with rotor rotational speed increase, the magnetic field penetration depth into the permanent magnet body and bandage reduces. Thus, the losses reduce either.

The article shows also that the magnetic system does not exert a significant effect on the eddy current losses, created by spatial harmonics, in permanent magnets. The eddy currents losses in permanent magnets herewith may alter significantly due to load angle variation.

The article individually considers the losses caused by temporal and space harmonics. It also presents their numerical evaluation and describes of their minimization techniques.

The article considers synchronous motors of a reversed structure with electromagnetic excitation of both conventional and based on high temperature superconductor tapes (HTST). The presence of excitation winding allows perform deep regulation, while current carry capabilities of modern HTS tapes of second generation allow create magnetomotive force (MMF) of the excitation winding, exceeding permanent magnets.

Synchronous electromechanical transducers of reversed structure with electromagnetic excitation open prospective application domains in wind-power engineering, low and middle power hydropower, special and military applications.

Based on analytical solution of the problems of magnetic fields distribution in active zone of electric motor of a reversed structure with both electromagnetic and permanent magnet (PM) excitation, the authors obtained equations for electromotive force (EMF) and inductive resistance. The obtained equations allow determine the dependence of motor's output parameters on the pairs of poles number, geometry of active zone, and excitation MMF. Likewise, in case of HTS tapes' implementation in the excitation winding (EW), it is possible to define the dependence of the motor parameters on the properties of the tape in use. Based on analytical equations the comparison of the motor of a reversed structure excitation MMF with PM excitation was performed. Besides, the analytical equation allowing compare these two types of excitation was obtained. It is shown, that the power of a motor with electromagnetic excitation can be greater with lower number of poles and HTS tape current close to 100 A. The obtained analytical equations can be employed for optimization calculations while defining the main sizes of a motor active zone. The combination of the presented fundamental solutions of theoretical problems and modern simulation methods will allow develop new calculation procedures for both traditional motors and motors based on HTS materials.

The all-electric (more electric) aircraft (MEA, РОА, МОЕТ) concept is currently the main trend in the development of the perspective aircraft power system both in the Russian Federation and abroad. This concept assumes the replacement of aircraft pneumatic and hydraulic actuators by electric (or electro-hydraulic) ones, as well as transmission generators' constant speed drives elimination.

The total rated capacity of MEA aircraft electric power supply system can reach up to 1,5 MW. To ensure the specified quality of the electric energy at the consumers inputs and mutual backup, its distributing channels should e connected in paralles. Thus, each channel should contain a higher voltage (270 V or 540 V) DC link in addition to the low voltage (27 V) central distribution unit with battery. These higher voltages are no used for feeding the distribution unit buses due to the complexity of arcless commutation provision. Thus, each MEA electric power supply channel is an independent combined AC-DC complex with four types of the central and peripheral distributing units: 115/200 V, 360... 800 Hz; 115/200 V, 400 Hz; ± 27 V and ± 270 (540) V. Electronic secondary power supplies interconnect these units with each other.

The authors suggest the structure of combined electric power complex with secondary power supplies' modular and scalable architecture for all-electric aircraft power supply systems with increased power-to-weight ratio based on unified multipurpose switched mode converters. This structure ensures parallel operation of both supply channels to improve electric energy quality.It reckons multiple mutual redundancy of the circuits for essential consumers feeding, and allows unify multifunctional switched mode converters, constituting it.

The considered above nonconventional MPC circuit solutions provide the required electric power quality in static and dynamic modes, high specific power (per unit mass and volume) and increased functional reliability. These solutions are protected by Russian Federation priority and provide high extent of import substitution in listed products of power electronics.

The spacecraft onboard equipment electronic components and units, as well as cable networks benchmark testing on electrostatic discharges (ESD) resistant strength should be carried out conditions closer to the real spacecraft operation conditions, in which electrostatic discharges occur.

Benchmark tests are performed in the air medium, and electrostatic discharges are simulated an ESD-generator. Thus, drawing near real conditions of the outer space is possible only by insulating elements and units from grounding circuit and maximum offset from the conducting environment to reduce the capacitive coupling.

Establishing the standard requirements to the onboard equipment noise immunity to simulated electrostatic discharges impact allows ensuring the possibility of comparative analysis of the testing results of various space vehicles.

These standard requirements should specify the simulated ESDs types; the degree of the tests' robustness; characteristics of the working place for tests. The testing methods should account for the specifics of onboard elements and units, as well as cable network placing on the spacecraft structure.

The article presents the description and requirements for the spacecraft onboard elements and units, as well as cable network benchmark testing.

The authors suggest performing the benchmark tests in such a way that the elements and units under testing together with along with spacecraft shell element and measuring equipment would not have connections with grounding circuits and power network, and placed far from the conducting medium.

Metallization of through holes in silicon wafers has been investigated by electron microscopy methods. Dependences of the thicknesses of metallization at various depths of through holes in wafers for single-sidedand double-sided sputtering of chromium and copper, with thickness of 1 µm to 5 µm, as well as with successive galvanic deposition of layers of chrome and copper and chrome and SnBi alloy (tin 98-99%, bismuth 1-2%) on the films of chromium and copper were obtained by vacuum magnetron sputtering method. Optimal modes of through holes metallization in silicon wafers process with closest characteristics of film deposition along o all structure elements, consisting in performing the process in two stages were determined. Initially, employing vacuum magnetro n sputtering method prepare metallization with minimum thickness, ensuring formation of continuous metal film. With two-sided metallization by vacuum magnetron sputtering of chrome and copper, the derived films have minimum thickness in the middle of the through holes. The continuous film is formed at chrome and copper thickness more than 1.7 µm on the surface of the through holes. To ensure the thicknesses it is necessary to perform two-sided sputtering of chrome and copper by vacuum magnetron sputtering methods with thickness less than 4 µm. Then, by galvanic precipitation method refilling to the desired thickness should be performed by galvanic precipitation method. Thickness changing at one-sided metallization sputtering, obtained in through holes by vacuum magnetron sputtering method presents linear decreasing character with increase of the holes' depth.

The minimum thickness of metallization is determined, at which a continuous metal film is provided along the entire depth of the through holes in the wafers. With a thickness of less than 1 µm, the surface of the film in the through holes is not continuous, but an island one. When sprayed from the front side, a continuous film forms on the surface of the plate, but the metal structure it is not continuous on the chamfers and walls.

This article subject of research are electrically conducting polymer compounds based on epoxy binder with non-covalently modified carbon nanotubes (CNTs). Such compounds can be applied as binders to create hybrid functional polymer composites. Fluoro-organo-silicon block copolymer was used as CNTs' modifier, which is organically compatible with epoxy olygomers. It allows regulate the interaction of the modifier with polymer matrix and study the nano-composite's functional properties under various distribution of the carbon tubes.

The goal of this research consists in developing a method to control electrical conductivity of polymer nano-composites by controlling the spatial distribution of CNTs in the bulk of the binder under development to create hybrid polymer composite materials with functional properties.

In the course of this work execution the experimental research on the development of a method of preparation of electrically conducting binder based on epoxy resin and non-covalently modified CNTs. Measurements of electric conductivity of hardened composition were performed. Since the non-uniformity of the CNTs' distribution over the nano-composite surface does not allow determine the value of the surface resistance with adequate accuracy by contact methods of conductivity measuring, the non-contact method was used based on measuring the electromagnetic wave reflection coefficient within the range of 20-35 GHz. The authors measured also the viscosity of the binder and determined the spatial distribution of nano-particles in the bulk of composition by scanning electronic microscopy and determination of element composition.

The effect of the modifier concentration on electrical conductivity and rheological properties of the binder was studied. It was established that the modifier concentration variation allows regulate electric conductivity of nano-composite and viscosity of modified binders under the constant concentration of CNTs. In the course of this work we obtained the values of electric conductivity of about 7.3 S/m with the viscosity of the developed binder comparable to the basic binder.

The results of the study allow solve technological problem of decreasing the viscosity of epoxy binder modified by carbon nano-particles, to produce electrically conducting hybrid polymer composite materials under conductivity preserving.

The lack of information in domestic and foreign sources on the effect of carbonaceous impurities purification technology on titanium alloys' properties complicates for technologists selection of the most effective and safe methods of purification of a gas turbine engine compressor air-gas channel parts and units.

The purpose of this work consists in evaluating the property change of VT20 titanium alloy while removing carbonaceous impurities by chemical means.

The studies were performed with laboratory samples manufactured from a VT20 alloy sheet-billet. Caronaceous impurities, imitating operational ones, were applied on a number of samples according to the developed technology.

Eight foreign made and domestic chemical technologies (compositions) were studied as purification means.

The authors established that the most effective removal of the carbonaceous impurities from the surface of the heat-proof VT20 titanium alloy was ensured by domestic purifying solution No 1, a two-stage purification technology in alkaline and acid solutions (“loosening + etching”), and foreign made solution HDL 202. However, while purifying carbonaceous impurities with HDL 202 solution a general etching of the surface and its microstructure change might occur.

The surface roughness values of the VT20 titanium alloy do not change significantly after the removal of impurities. The relief and profiles of the purified surfaces have a shape similar to those of the original samples.

A slight increase in the microhardness of the purified samples (up to 5%) can occur due to gas saturation of thin surface layers, due to both formation of carbonaceous impurities and the processes of chemical surface purifying.

When purifying the surface from carbonaceous impurities, the activity of the surface decreases, regardless of the type of the solution used. The least decrease in activity is ensured by cleaning solution No 1.

There is no deterioration of moistening characteristics by the VPr16 solder of the surface purified from the carbonaceous impurities by purifying solution No 1 or two-stage “loosening + etching” technology and HDL 202 solution.

Purification of carbonaceous impurities by all studied solutions does not lead to VT20 alloy strength and plastic characteristics degradation, and to a change in the character of its destruction under conditions of static loading.

The article presents specifics of orders formation depending on their basic characteristics.

The presented work discloses the fact that at present the system of orders formation and realization process is of paramount importance. This is important so that the orders themselves should be dealt with in the context of their life cycle. The orders can be split between each other according to such parameters as volume, liability distribution while their fulfilment, delivery periods and formation method. The process of the order commissioning or the order lifecycle can be conditionally split into certain stages. The life cycle stages would differ from each other for various types of the order commissioning. Therefore, one should have an idea of an order life cycle specifics for various types to minimize the error occurrence probability in strategy selection, and, as a result, minimize the probability of financial losses risk occurrence.

The article consists of three main parts, namely, introduction, the gist of the work and conclusions. It presents the description of various stages specifics, and determines the relationship of the enterprises activities marketing component and the logistic one. An algorithm for the order maintenance was developed, and efficiency evaluation technique is presented for one of the order types.

Finally, the authors describe the orders formation technique through the Internet media and its efficiency evaluation tool. The article presents the basic technologies for the delivery of the order, best suited to the market demands.

Eventually, it is worth noting that companies should master orders fulfilment and maintenance methods through Internet medium, since at present this is the prospective trend of development. Besides, these methods are fast and economically sound as well.

When solving the problem of objects' multicriterion selection and seriation, modern mathematical modeling technologies can employ various methods, including simple aggregate weighing (SAW) and “ideal point” (TOPSIS) methods.

When comparing alternative objects of research by their effectiveness, the necessity occurs to account for not only positive or negative indicators, but estimate the objects by the degree of proximity to a specified value, i. e. to the criterion. It should be noted, that the criterion value could lay in the middle of the range of the considered indicators. Besides, the object's effectiveness, as a rule, has non-linear dependency from the change of the indicator value. Inasmuch as the existing algorithms of SAW and TOPSIS methods do not allow perform such task, a certain modernization of the TOPSIS method is required. This method is top-of-the-line with respect to the ranking procedure.

In general, when the case in hand is the indicator's degree of proximity to the specified value, the attainability function is used. In multicriterion analysis, it is called the utility function. It allows realize transformation of the initial “decision matrix” system into normalized matrix, with account for the proximity to the specified values of the criteria. This operation is close in its meaning to linear (nonlinear) normalization. It is performed in the SAW method (determines the degree of the maximum or minimum attaining), and replaces the rationing using the TOPSIS method.

Earlier, the TOPSIS method could be applied, when a monotonic-increasing utility function existed for each criterion. In other cases, one had to apply the more simplified SAW method.

The presented TOPSIS method modernization gives, firstly, practically a comprehensive agreement of the computational results with the above said methods for positive indicators, and, secondly, a slight difference with the SAW method while using positive and negative indicators, when the unknown function of the relationship between efficiency and indicators is non-linear (linear and non-linear normalizing).

Thus, the proposed modernized version of TOPSIS method allows extend the scope of this method in the case of the specified criteria values (positive and negative), located within or outside the range of the indicators variation.

Startups become the benchmarks of innovation growth, and the government is interested in their successful functioning. It is proved by the formation of support infrastructure around startups. An integrated approach is required to startups' support realization.

The article describes the life cycle model of a startup. Irrespective of the startup's type and line of activity, each of them passes typical stages in the course of its development. The startup lifecycle model is associated with the I. Adizez's model of the organization lifecycle.

In the beginning of their functioning, most startups face with the necessity of solving the similar problems and performing the similar functions, particularly with the necessity of the substantiated selection of the way of their development.

The management mechanism of industrial startup development allows obtain a justified choice of the way for its further development. The participants of an expert group obtained a conclusion on the most effective method by carrying out the procedure of a startup development evaluation. However, it is necessary to ensure the effectiveness of the taken decision.

The expediency of applying the mechanism and the efficiency of the taken decision should be confirmed by calculations of economic efficiency. Thus, the methodological approach to the management mechanism of the industrial startup development becomes a crucial issue. All approaches consider the object of evaluation from certain sides, and are based on specific external and internal information.

The article describes the approach to efficiency determination of the industrial startup development mechanism management. A structure for calculating the total costs of a startup during the period of itscommercialization with the chosen development process was developed. A typical process for analyzing and calculating the total costs of a startup is presented.

Monitoring of the current state of the aircraft fleet is an essential component of the management process of the aircraft fleet renewal by introduction of new types of aircraft with improved technical and economic characteristics.

The branch (Federal Air Transport Agency -Rosaviatsya) does not supervise (keeps record) on such deals as purchase and leasing of foreign made aircraft operated in Russian Federation. We mean the monitoring with respect to concrete transaction number, the date of leasing commencement and its expiration.

In this connection, the problem of such aircraft retirement from the Russian air transportation market is difficultly formalizable. This analysis was based on publicly available databases Flightglobal (http:// dashboard.flightglobal.com/app/fleet/#/analyser/fleet), contract data and published reports on big deals of the companies.

For the purposes of the analysis the operated aircraft was separated into several groups according to ownership types and aircraft age. These groups are as follows: the aircraft owned by air carrier; aircraft obtained by financial leasing (from which the author separated out the subgroups of “young” aircraft, the aircraft with life span lower than 12 years, and the aircraft with life span more than 12 years); the aircraft in back leasing and operational leasing.

The owned aircraft retirement was determined according to the expected life span, or maximum permissible flying hours and endurance cycles from the commencement of operation.

The retirement of an aircraft being in financial leasing or leaseback can be forecasted only by life the span and total operating time.

Meanwhile the economic mechanism of the transition of the second-hand aircraft from the big companies to regional Russian companies is not developed.

The article presents some results covering the general situation in passenger aircraft fleet in the branch at large, and more detailed on Aeroflot Russian Airlines.

At present, the problems of sustainable development of socio-economic systems of various levels are in the sphere of close attention of Russian and foreign scientists.

The goal of the article consists in considering the possibilities of sustainable development of the northern territories, which have pronounced specific features. To clarify the concept of sustainable development of the northern territories, the article reveals limitations in the use of their resource potential, namely: high level of production costs, high vulnerability of the natural environment, preservation of the traditional way of life and the need to improve the quality of life of indigenous small peoples.

Based on the revealed limitations, the article formulates the principles on which the sustainable social and economic development of the northern territories should be based, the priority role of which is assigned to the implementation of the latest technologies to ensure high quality of life for the population and improve the ecological situation.

Remoteness of the Northern Territories from large settlements, its difficult climatic conditions, and the poor development of land infrastructure predetermine the high importance of air communication for the sustainable social and economic development of the North. Thus, the latest technologies used for the development of the Northern Territories can be concentrated in this area.

The article analyzes the experience of Canada and Alaska in the development of air communication. These areas were selected for analysis due to fact that natural and climatic conditions, labor endowment provision level, the distance from the economically developed regions allow establish certain similarity of these territories with the territories of the North of Russia.

Based on the performed analysis of international experience, the article suggests the following opportunities for using modern aviation and space technologies to ensure sustainable social and economic development of northern territories:

– Development of hub airports, as well as regional and local air transportation, supposing inclusion of regional and local airports into nodical structure of air transportation servicing;

– Implementation of unmanned aircraft for the delivery of goods;

– Implementation of high- capacity space communication vehicles to create a broadband satellite communication network;

– Remote sensing of the earth surface for natural resources development while meeting the requirement to preserve the unique ecosystem of the North.

The considered technologies are quite expensive and require significant investments in research and development. In this regard, the opportunities for sustainable socio-economic development of the northern territories are closely associated with the role of the state in socio-economic development.

The article concludes that the formation of an effective system of interaction of federal and regional authorities, business and indigenous people, the optimal combination of market mechanisms and public administration tools will ensure the implementation of opportunities for sustainable social and economic development of the northern territories and level the specifics of these territories that complicate this process.

Ribbed panels from aluminum alloys are widely used in aircraft industry as power structural elements, parts of the wing and fuel tanks, as well as in the form of the heat-exchange surfaces. Increased requirements on strength and reliability are rendered to such panels. The most rational technique for such kind of panel manufacturing, i. e. isothermal punching, may lead to clipping in the ribs and sink marks on the side, opposite to the ribbing.

Modeling and experimental results of the studies reveal that defects stems from the combination of manufacturing process control parameters, such as temperature and deformation velocity, as well as geometrics, i. e. blank thickness.

The main objective of the studies consists in developing design technique of the part blank design as a function of temperature and deformation velocity while isothermal punching.

The put forward problem is solved by control polynom development, linking manufacturing process parameters – the blank temperature, velocity and geometrics with the defect magnitude, i. e. sink marks in the ribbed aluminum panel while its manufacturing by isothermal punching technique.

The initial data for the required polynomial is the results of finite element mathematical modeling with varying initial parameters of the punching process and the magnitude of the forming sink mark or its absence.

The obtained modeling results were processed according to the three-way analysis of variance planning procedure. The regression equation was obtained to compute the sink mark magnitude in the ribbed panel in dependence of the process temperature and velocity, as well as the initial blank thickness.

The authors applied the analysis of variance, which allowed define the significant factors in the calculated polynomial, and, neglecting the rest, significantly simplify it.

The sink mark magnitude obtained with the calculated polynomial correlated well with the results of mathematical modeling and experimental studies.

The proposed method is universal and can be implemented for various cases of defect-free technological processes design, when evaluating the impact of the process's control parameters on and their contribution to the manufactured product's characteristic being studied is required.

Presently, one of the meaningful problems in modern machine building consists in creating new electric power sources. It is of special importance for such human field of activities as aircraft and spacecraft building.

Chemical current sources (CCS) based on aluminum anodes, where various solid, liquid or gas oxidizers used as cathodes, found an extensive application in spacecraft electric power systems. Very often, the modern methods of such sources design, however, turn out rather complicated and energy consuming.

One of the most successful electrical systems based on CCS, applied in modern spacecraft, is an oxygen-aluminum system with liquid electrolyte, consuming oxygen from the environment.

In most cases, anodes are made of aluminum alloying with such metals as Ga, Sn, In. Nevertheless, the high values of anode potential and current were obtained while the experiments with Al-In alloy anode. It was found, that aluminum doping with Indium ensures anode electrochemical activity with faraday efficiency no less than 90%.

Thus, this work was focused on developing the scientifically substantiated technology of anode manufacturing based on Al-In alloy to ensure highly dispersed, isotropic structure to provide a uniform anode dissolution, decreasing pitting formation, and, as a consequence, increasing energy and performance characteristics.

The main difficulty while Al-In alloys casting consists in organizing a uniform indium particles distribution (which is not soluble in a solid aluminum) over the solid base metal volume. The reason to it stems from too large difference between aluminum and indium melting points (659°C and 156°C respectively), as well as high density of the latter (6.5 g/cm³). Introduction of traditional modifiers into the alloy is unacceptable, since they (Ti, Zr, B) aggravate the electrochemical figures.

The studies conducted in MAI (laboratory UNPL “TOMD”) allowed develop technological scheme for obtaining anodes' blanks. The scheme includes obtaining ring blanks using additive technology of centrifugal casting and pressing by using shear deformation during pipe billet extrusion. It will allow work out sufficiently the alloy structure, grind up the phase inclusions and, as a result, ensure the necessary properties' level.

A distinctive advantage of the developed technology compared to the analogues is the possibility of regulating of a significant number of factors during the deformation process and, accordingly, to obtain the best possible material's characteristics in the final product. This technology is realized herewith using traditional equipment.

The presented work deals with considering pneumo-thermal molding and diffusion welding (PTM/DW) technology for multilayer structures manufacturing from titanium alloys, including the ones of variable height. The paper represents the presentation of general theory of the above said technology, and analysis of the problems emerging while its realization.

The author separated out the stages of PTM/DW technology of multilayer titanium panels.

The main problem considered in the paper consists in the problem of non-removable defects formation, accompanying manufacturing of multilayer wedge-like panels. These defects are imaged on the appended plots and figures.

The reason of these defects occurrence while multilayer panels molding lies in the different displacement of the lower shell in various areas of the pack of sheets. In the area of diffusion welding this displacement is constrained by ribs of the filler, while in the zones which are not welded with the filler, the upper shell is forming freely under gas pressure. Its deflections are forming herewith between the areas of welding with the filler.

Solution of this problem consists in defining the managing program, necessary to form the ribs of the filler and the shell, whereby the shell deviation in the areas unreinforced by the ribs would not reach critical value.

The recommended range of the shell and filler thicknesses ratio in dependence on the shell deflection in the areas unreinforced by the ribs, as well as equations for determining critical deflection factor and molding pressure were obtained by mathematical modeling.

Application of the above said equations for the filler and shell thicknesses of multilayer wedge-like panels will allow avoiding defects occurrence, which was confirmed in practice. All fabricated panels comply with the calculated parameters. No defects were detected over the profile section dimensions.

In view of the foregoing, one may state that the problem consisting in determining the regularities for selection of design and geometry parameters of multilayer structures, allowing ensure qualitative molding process without defects formation was solved successfully, and the solution has practical applicability.

The main objective of the research is the direct integration of the electric machine such as generator or starter-generator on the low-pressure and/or high-pressure shaft of an aircraft engine, and the gearbox elimination. This will allow reduce the aircraft engine's weight and size figures, as well as improve the aerodynamic efficiency of an aircraft as a whole. This article presents the design and experimental research of the scalable prototype of high-temperature starter-generator with the inner rotor for more electric aircraft. The fundamental difference of the developed generator from the conventional machine consists in no oil ingress into the rotor or stator cavity. The starter-generator is immersed in an aircraft engine oil chamber, containing the oil necessary for bearings lubrication at the temperature of 120-160 °С. The stator and rotor are not lubricated with oil, which does not circulate. Cooling is achieved by losses' heat sink into the surrounding oil. A scalable high-temperature starter-generator prototype model was developed in Ansys Maxwell software package. It revealed a high accuracy and close convergence with the experimental results. Moreover, the system efficiency assessment and computation of losses in starter-generator's elements were performed. Based on the experimental results and computer simulation the starter-generator full-sized model was developed, and tests at the temperature of 120 °С were conducted. This generator appeared to be less loaded from the viewpoint of electromagnetic and thermal loads. It proves the efficiency of the proposed conception and its effectiveness for implementation in more electric aircraft.

The paper compares competitiveness of Il-76MD-90А with the US C-17 military transport aircraft. The basic assessment criterion is the aircraft capability to perform landing on unprepared sites with restricted run length, which requires employing the engine thrust reverse.

The problems under discussion relate to employing thrust reverser of PS-90A-76 engine, installed on the Il-76MD-90А aircraft. These problems do not only increase the cost of an aircraft operating cycle and affect the flight safety, but reduce its competitiveness as well. The paper presents computation and experimental data, revealing that the main cause of the emerging problems consists in poor external aerodynamics of the power plant during an aircraft ground run employing the thrust reverse. By external aerodynamics the authors mean the gas jet discharge type form the engine reverse units, which may interact with the engine itself and control airframe surfaces while its ground run. Such interaction can lead to:

– gas dynamic instability in engine operation;

– damages to the rotor blades of the engine caused by foreign objects thrown from the surface of the aerodrome;

– Aircraft dynamic characteristics deterioration (wind drag, stability, controllability), and aircraft run-length increase. Unsatisfactory external aerodynamics of the Il-76MD-90A aircraft is the cause of its poor competitiveness compared to the US military transport aircraft S-17.

Ways to the aircraft external aerodynamics improvement are considered below:

– the engine reversal thrust value optimization;

– reverse jets discharge optimization in accordance with the aircraft layout.

Ways of the Il-76MD-90A aircraft external aerodynamics improvement were developed based on estimated and full-scale studies. The substantiation of the developed measures is based on the design features analysis of the S-17 engine reverser unit.

From the above said the author concludes:

1. The level of Il-76MD-90A aircraft power plant external aerodynamics is not high enough.

2. The Il-76MD-90A airplane competitiveness recovery requires carrying out studies on the power plant external aerodynamics improvement, which will allow competition with the S-17 aircraft in the foreign market.

Development of new up-to-date actuating gas turbine engines based on fourth generation aircraft engines requires certain time consuming. Thus, one of the ways to series-production engines' parameters improving consists in their upgrading and forcing. A fifth series-production NK-16-18ST engine was developed hereupon at SC KMPO. Its more productive high-pressure compressor allowed ensure higher flow velocity (about 170 m/s) at the combustion chamber inlet. Combustion organization and provision of optimal level of toxic agents' emission in engines of such kind is hindered due to high-pressure parameters of the airflow.

Such situation led to the necessity for carrying out research and design effort consisting in altering structures of burner and flame tube with redistribution of air vents along its length. The approach, used in the above said structure lies in forming the «reach» mixture in combustion chamber primary zone with its subsequent sharp weakening to ensure “poor” content, which allows maintain low level of nitrogen oxides. Testing of this chamber together with the engine confirmed that the sel ected approach allows reduce nitrogen oxide content in combustion products. However, it requires a number of measures related to the structure changes to achieve the desired level of noxious substances emission. To increase penetration depth of a jet into combustion zone the chamber was upgraded by cylindrical hubs installation in the first row of vents. This measure allowed reduce concentration of oxide nitrogen emission in the engine's exhaust gases, but it was not enough to ensure the level, required by regulations. Having in mind, that residence time reduction of gas in high-temperature zone decreases oxide nitrogen formation, in the framework of the last version, measures were introduced to increase fuel mixture flow velocity fr om atomizers, mounted on the flame tube head. The atomizers have elongated nozzles and less diameter. Such configuration allowed ensure noxious substances emission in combustion products at the level complied with the State Standard requirements (GOST 28775-90).

The informationally and navigationally oriented spacecraft injection to the working orbit with high positioning accuracy, scientific and research spacecraft transition from support orbits to departure trajectories for deep space flight and other complex tasks of space exploration are carried out by rocket transportation systems. These systems include specialized withdrawal means, named “upper stages”. The following requirements, such as enhanced energy efficiency and reliability, long-term staying in starting readiness mode, protracted operating time and multiple starts are imposed on upper stages' cruise engines. The «liquid oxygen-liquid hydrogen» cryogenic pair burning engines possess maximum energy efficiency. The first home-produced oxygen-hydrogen LRE is 11D56 engine developed at Khimmash Design Bureau headed by A.M. Isaev. This engine can be considered as the basic one for ecologically clean upper stages for rocket carriers of “Angara”, “Soyus 2-16” and “Soyus 3” families presently under development. This engine's design allows modernization or modification (without significant time consumption) of its structurally stand-alone units, preserving characteristics, which define the engine workability and reliability at large. The KVD1 engine energy parameters and characteristics updating is realized by structural scheme optimization based on the structure technical analysis and effective options selection, related to the engine usage tasks.

Based on the experience in the KVD1 engine chamber design and development two options for chamber with retractable nozzle headers design were considered. For these options, corresponding to the two engine modernization variants, optimization of nozzle divergence geometric degree was carried out. Calculation of working process parameters and the main chamber characteristics optimization was performed.

The specific impulse's increase is analyzed by optimum relationship selection of fuel components consumption and selection of the maximum (optimal) pressure in the combustion chamber selection. The optimality criterion of fuel components consumption is payload weight maximum at geostationary orbit, at which, according to the specific impulse mass equivalent, the mass gain is equal to the fuel tanks of the engine unit the mass gain. the results of theoretical and calculating studies consists in defining principal design solutions of two variants of oxygen-hydrogen engines' chambers, under development based on KVD1 LRE.

“Green fuel” is an aqueous solution of a high-energy oxidizer (hydroxylammonium nitrate and others), and a fuel, presented by various substances, such as alcohols, glycerin, etc. It offers a number of advantages, namely, a higher density, low freezing temperature and high specific characteristics. Such mixtures relate to low-toxic substances, whereas hydrazine is a high-toxic substance. Thus, the “green fuel” mixture implementation as a monopropellant for an aircraft correction and orientation thermocathalytic thrusters is up-to-date issue.

Hydroxylammonium nitrate was sel ected as a basis for the “green fuel”, to which a fuel and dissolvent (water) are added in calculated ratio. The energetic qualities of the fuel depend on its basis, though its output characteristics are strongly affected by the water content in the mixture.

The laboratory model consists of a heater for the structure's starting warm-up of the, combustion chamber fr om refractory metal with a special protective coating, catalytic bed, consisted of a combination of metallic and granulated catalysts, an injector unit ensuring operating pressure differential, and a system of thermal screens.

The laboratory model presents a disassembling model to monitor separate elements of the structure. Besides, such model allows quick replacement of the thruster elements, such as the catalyst bed.

The laboratory model was tested in air under normal climatic conditions. The thruster was tested on firing functioning both in impulse and continuous operating modes.

The tests of the thrusters were conducted in continuous modes at the inlet's dropping pressure. It is worth mentioning that with the inlet's pressure decrease, the pressure in the combustion chamber decreases proportionally, which demonstrates the stability of the thruster operation.

The K100E laboratory model maximum run amounted to 1.5 kg of consumed fuel over 1500 start-ups. The main reason for the thruster's failure relates to the tests conduction in atmospheric conditions, namely to the oxidizing and destruction of separate parts of the laboratory model (heater, screens) under higher operating temperatures.

The paper tackles the version of the mobile system for fixing and accounting for aircraft damages while preflight checkup. The benefits of the system are defined by the possibility to use mobile devices of a tablet type equipped with high-resolution cameras. These devices fix the detected damages, and convey the images of these damages into the server part of the system, which performs their processing and automatic logging to various exploitative documents, related to aircraft servicing. The basic tasks resolved by the developed and described in this paper bundled software of the mobile system are formulated as follows:

The developed mobile system's software solves all above listed tasks, which allows accelerate not only the preflight checkup itself and filling out the related documents, but also the subsequent technical servicing (such as repair, gathering of statistical reporting systematization on the condition of the certain aircraft

1. Realization of the possibility to use mobile hardware (a tablet with high-resolution camera, or special hardware) for damage registration and the possibility to examine the tolerances on such damages by its location.

2. Solving the problem of a picture of damage operative “attachment” to the “structure damages list”.

3. Realization of automatic documentation generation to send the request for the element repair to the aircraft manufacturer, if such repair is beyond the scope of the Repair manual.

4. Loading to the database the information on the performed repair and substituting of the damaged elements picture for the new one.

5. Return the list of existing damages of the aircraft on the request of the program.

6. Performing checkup and repair according to instruction manuals. Furthermore, the program should provide an opportunity for regulatory documents viewing to determine further activities for the damage elimination.

One of the main factors affecting reliability of aircraft articles is vibration effect during the joint flight with carrier. To obtain estimates of reliability characteristics flight test of products are carried out. Modern equipment for ground vibration testing, reproducing the flight conditions, allows substantially reduce the amount of flight tests by replacing them with laboratory tests. The actual problem here is formation of laboratory tests regimes to ensure the equivalence of loading in laboratory conditions and in flight. Characteristics of vibration loads are obtained usually based on measuring data obtained during flight tests of the product or its prototype. At vibration tests, a relation is established between laboratory test modes and flight dynamic loads by the levels of vibration accelerations or stresses.

The paper presents the technique and results of flight and laboratory vibration tests on definition of vibration stresses and accelerations characteristics of an aircraft product's structure in typical flight. Laboratory tests were conducted with random dynamic loading, corresponding to loading during flight structural tests. The purpose of laboratory tests is determination of characteristics of a structure's accelerations and stresses in the conditions of a spacecraft joint flight with the carrier. This requires reproduction of conditions exhibiting adequately enough the loading condition of a product according to the basic probabilistic characteristics during typical flight.

The authors developed the technique and modes of aircraft product's vibration tests, complying with vibration loading of a product at every stage of the host aircraft flight. Using the obtained modes the tests were conducted, whereby the random dynamic loading, corresponding to the operation conditions of the product on an internal suspension bracket of the carrier, was realized. Comparison of vibration acceleration probabilistic characteristics at laboratory and flight tests demonstrated conformity of these tests' results according to root mean square values of vibration acceleration.

The developed laboratory tests technique ensures correct reproduction of random vibration loading reproduction of and aircraft product structure during the flight on an internal suspension bracket of the carrier. The technique and results of the tests can be applied for estimation the structure vibration strength of an aircraft product of various applications during the joint flight of an aircraft product with the carrier.

At present, the onboard aircraft defense structures for protection from “air-to-air” missiles are equipped with the systems of jamming cartridges of various calibers ejection. The existing algorithm of airborne defense systems application consists in practically continuous ejection of a series of jamming cartridges when the aircraft enters the area of the enemy's air defense. However, the existing techniques of jamming cartridges implementation do not ensure the aircraft protection from the missiles equipped with matrix photodetectors. There is a contradiction between the potential onboard defense systems implementation efficacy, and military characteristics of existing onboard defense systems. In this paper, the authors propose a technique for guided missile coordinates determination errors to ensure its neutralization on the flight trajectory.

A methodology for probability estimate of aircraft skipping the hit by a guided missile while realizing the active element ejection to the trajectory of the guide missile with its subsequent detonation was developed. The probability estimate of guided missile missing its target is based on probability calculation of the active element's detonation coordinates center will appear inside the dispersion ellipse with the main axes equal to mean square deviation of the active protection element detonation coordinates from the actual position of the missile. The equations for the active protection element detonation coordinates were obtained using measuring errors theory methods on the assumption of the aircraft and missile rectilinear and steady motion, after the initial miss elaboration, with an allowance for the aircraft and missile coordinates and flight speeds measuring error, as well as current missile's angle of attack determination errors. The paper shows that the missile's angle of attack determination errors depend on aircraft and missile current speeds determination errors, as well as missile bearing in vertical plane measured by optical radar station or specialized onboard radar station belonged to aircraft protection structure.

At present, methods for orbit parameters changing with cable system by periodical changing of its length for the case, when the cable is located in the plane orthogonal to the orbit plane, or when it lays in the orbit plane, are known. However, the orbit parameters changing is possible only in case of non-central gravity field. The presented paper offers the structure of interorbital transfer of the space system, consisting of two masses repulsed and retracted in the orbit plane by periodically formed coupling. The flight is considered in the central gravitational field.

Originally, the system represents a single spacecraft, consisting of two parts of equal weight. Initially, the system s on a circular orbit. The mass repulsion occurs in the direction tangent to the trajectory. Hereafter, the masses being uncoupled move independently of one another over various trajectories. Performing a various number of turns around the attracting center, after a certain period of time they will turn up on the line coinciding with the radius vector. One of the masses herewith will pass the pericentre of its orbit, while the other – its apocentre. At this moment, the masses contraction occurs assisted by the formed coupling. Methods of coupling formation are not considered in this paper. The paper demonstrates that the eventually formed orbit differs from the original one.

The authors obtained analytically the dependence of the system final velocity in the point of masses contraction after their contraction versus the speed of their repulsion ΔV.

The dependence of the masses contraction point radius vector versus the initial repulsion speed ΔV for the final orbit.

Here:

Solution of this problem revealed a theoretical possibility of orbit parameters changing for the system of a proposed type.

The analytical dependence of the speed value at the time of contraction versus the initial masses repulsion velocity is obtained.

The equation determining the radius vector in the point of masses contraction of the formed final orbit created.

Active development and application of digital technologies enabled realization of advanced algorithms in aircraft control systems, which could not be implemented earlier due to the limited capabilities of analog computers, and the more, so in mechanical control systems. The attempts to ensure aircraft control characteristics invariance to varying flight conditions, aerodynamic configurations, centering and mass-inertia characteristics led to the necessity of employing two classes of characteristics onboard the aircraft, namely, with reference model and with the identifier. Such algorithms are developed and successfully applied in the control systems' longitudinal channel of Su-30Sm, Jak-130, Su-35 and Т-50 aircraft. It is important to notice that the adaptive algorithms in lateral control channel have not found practical application, though the problems requiring solution exist there either.

Thus, the problem of lateral controllability deterioration caused by occurrence of the adverse moment in yaw during aileron deflection exists on all MiG-29 modifications. The aircraft heel moment caused by lateral static stability due to the sliding is directed opposite to the effective aileron roll control moment. Flight speed reduction and increasing angle of attack corresponding to it lead to reduction of available rate of roll, and in limit case to occurrence of roll back reaction to control stick deflection. The acuteness of the problem is partially reduced due to implementation of a unique structural solution, i. e. airspeed head wind eddies generators, which, however, does not eliminate the problem at large. The situation is aggravated in the case of external suspension brackets asymmetrical mounting, which becomes a standard situation in view of the increasing effectiveness of aircraft means of destruction. The aircraft herewith begins to react differently to the stick deflection to the suspension bracket side and to the side, opposite to it, by the rate of roll. It complicates substantially delivering air combat and ground targets attacking, accompanied by drastic banking maneuvers of positive sign to negative and vice versa. Aircraft landing approach control is rather complicated, especially in case of landing on a ship deck of limited size in conditions of oscillatory motion and atmospheric turbulence.

One of the trends of gas turbine engines perfecting consists in “intelligent” engine developing. Within its control system, a so-called “virtual engine” functions in real time mode, i. e. a full range thermo-gas-dynamic GTD model. Its implementation allows, in particular, realize engine control by its critical parameters inaccessible for measuring. The gas temperature in the hottest part of the engine duct, i. e. the temperature at the turbine inlet, is one of such parameters. The paper presents the result of the study of new turbine cooling control methods, differing fundamentally from conventional indirect open-loop control of air bleed valves according to rotation speed, employed in modern automatic control system. A block diagram and algorithmic provision of adaptive closed loop control of turbine cooling units operating directly according to gas temperature prior to the turbine and rate change of turbine blade temperature are considered.

The result of such type of control estimation, carried out as applied to modern turbofan engine with high bypass ratio, revealed that its' implementation may allow:

– Engine efficiency increasing by decreasing the bleeding air consumption;

– Engine lifetime increasing by turbine inlet temperature decreasing by 100...200 К at steady-state modes, and the rate of turbine blade temperature decreasing by more than 20% at transient modes.

An adaptive control of air bleeding for turbines cooling associated with gas temperature limitation by effecting on the fuel flow in the combustion chamber was considered. The paper demonstrates that its implementation is possible:

– In flying conditions, when maximum engine thrust is required. It can be increased by 10% with the maximum allowable limitation of turbine blade temperature;

– Under operation conditions when engine lifetime is critical. It allows blades temperature reduction by approximately 50 K while maintaining the thrust value and specific fuel consumption.

The paper addresses the multi-purpose mathematical model of the electro-mechanical actuator's (EMA) dynamics. It contains the general description of the EMA, which was the object for the modelling, the description of the mathematical model developed and mathematical modeling results. The actuator was developed in the frame of the Russian-European project called RESEARCH for the elevator deflection of a regional passenger airplane. The mathematical model was implemented within MATLAB/Simulink software.

The actuator model consists of four submodels of its physical constituent parts such as controller, power electronics block, electric motor and mechanical gearbox (ball screw transducer). Programmatically switchable models with various level of detail of physical processes were realized for each part. The electrics were realized by the submodels of a single-phase DC motor and a simplified controller corresponding to it. It also contains three-phase induction motor with permanent magnets, regulated by a controller, realizing vector control in {p, q}-coordinates. Power electronics is modeled either by simplified dynamic elements, or on a physical level in detail (electronic components level). Special attention was payed to mechanical part of the actuator modeling, i. e. various submodels of non-linear mechanical effects of a “dry friction” and “backlash” were realized. Thus, we managed to ensure a balance between modeling accuracy and speed within the framework of a single model.

Based on mathematical modeling results the paper demonstrates how the dry friction and backlash parameters, as well as software methods of their realization effect on the actuator's regulation quality and its characteristics. It shows that program splitting of the actuator states (idle, motion, initiation) based on velocity smallness without using the sign function approximation is optimal method of dry friction effect accounting. It ensures reproduction of the necessary actuator motion pattern with acceptable integration step (10-4 s). The paper demonstrates also that accounting for linear stiffness of the actuator's ball screw transducer has insignificant effect on the actuator's frequency response within the frequency range of control surface control. It is shown that the replacement of the three-phase motor with a single-phase one while reducing the EMA model leads to different regulation character even while using the similar regulator structure and comparable PID-regulator coefficients.

The developed model can be used while the electromechanical flight control systems design for various engineering tasks, characterized by significantly varyng requirements imposed on the model in use. These tasks include: 1) development of the actuator and its control system, including actuator digital regulator synthesis; 2) actuator static and dynamic characteristics express-analysis; 3) obtaining reference actuator characteristics including small control signals; 4) analysis of transient responses and stability margins of the closed “aircraft – flight control system – actuator” control loop, including in-line simulation; 5) study and optimization of actuator thermal conditions while operating in the closed bay of the outer wing.

According to the item of the State Program “Development of the aviation industry for 2013–2025” creation of scientific and technological capacity ensuring global leadership in aviation technology and product promotion of domestic aviation industry on the domestic and foreign markets should ensure high competitiveness of domestic aircraft by introduction of innovative developments. Operation and maintenance analysis of the existing civil aircraft park revealed that while aircraft aerodrome maneuvering the hundreds of kilograms of kerosene are wasted, and drive trucks waste tens of kilograms of fuel. When kerosene burns in an aircraft engine, and fuel burns in combustion engine the atmosphere is contaminated by noxious substances. While aircraft maneuvering on the runway the noise level is 90 dB. Using combined actuator in landing gearwheels will allow decrease negative effect on the environment and eliminate completely the majority of shortcomings.

The paper presents the schematic diagram of electromechanical landing gear wheel actuator in which brushless switched-reluctance motors are mounted inside cylindrical gearwheels. Due to low cost materials implementation, small size and weight, low energy consumption and high efficiency maintainability better design and operating characteristics of aircraft landing gear wheel actuator are ensured. While motor-reducer design, specifics of its operation in the landing gear wheel were accounted for. The results of motor-reducer computation, which demonstrated the wide specter of implementation of such kind of actuator for various types of aircraft components, such as landing gear wheels actuators, high-lift devices' elements are given. The presented motor-reducer possesses diversified structural concepts, which allows use it for various types of aircraft both civil and military oriented, as well as for unmanned aerial vehicles (UAV) and spacecraft of various kinds. The prototype of motor-reducer, used for UAV's high-lift devices, displayed its apparent advantage compared to the other actuators, such as design compactability, manufacturability and cost effectiveness. Implementation of the above-described structure will allow fuel consumption saving by both an aircraft, and airfield servicing facilities. The structural concept of the motor-reducer in aircraft landing gear wheel does not have counterparts either in Russia or abroad.

The paper offers new technological solutions for gas turbine engines (GTE) manufacturing. These solutions are based on special processing methods using and additive technology implementation. To improve technological process of GTE parts manufacturing the authors suggest new technology of polymeric tool electrode (TE) fabrication with subsequent metal coating of its work surface using electrodeposition method. The most complex problem consists in ensuring accuracy of the profiled surface in the process of electrodeposition of a metal layer.

The objective of the work is developing computer model of the process of electrodeposition of metal on a dielectric TE for electrochemical machining (ECM).

This presented method consists in creating the information model, and studying the main process parameters of electrochemical deposition: electrolyte and electrode surface potentials, electrode reaction behavior, thickness and uniformity of the coating. While analytic model development parameters of electrode reaction, such as the exchange current density; electrochemical anodic and cathodic transfer coefficients; system electrode reaction equilibrium potential were determined. Besides, the above-mentioned method includes comparison of the formed profile with theoretical one.

The developed information model demonstrates that the metal coating possesses a variable thickness. On the boundary of cathode with electrode junction a thickening, stipulated by electrochemical processes, was formed. Here, in this zone the thickness deviation of the formed profile from a theoretical one is 355 µm. This implies that a minor mechanical processing is needed.

The developed technology allows carry out technological regimes and using the ECM obtain more detailed information on surface shaping.

To ensure the specified shielding efficiency electromagnetic screen should be homogenous to the maximum.

The uniformity of the shield depends on the resistance between the cable shield, the electrical connector and the onboard device case, i. e. transition resistances. High shielding efficiency can be ensured with small values of transition resistances.

The transition resistance is not a constant and can variable significantly during the life cycle of a product. These variations are caused by the effect of various factors: shields and cases bonding and connecting techniques; temperature and environmental conditions; operating conditions.

The results of the experiment that simulating a stay in a tropical climate revealed, that the magnitude of the transition resistance has increased up to 8 mOm, and in some cases it increased from 1 to 26 mOm, which significantly exceeds the standard value.

While temperature fluctuations effect testing, cable connectors subjected to thermal shock by immersion in liquid nitrogen with subsequent heating to 290°C by the stream of hot air. The results of this experiment demonstrate, that the transition resistance of the heated connector increases from 1 to 6 mOm.

In all these experiments, significant changes of transition resistances values in the direction to increase without returning to the initial values were observed. The reason for this consists in the thermal deformation of the parts' shape and contact failure due to the emergence of the oxidized layer.

The results of the shielding effectiveness study show that the magnitude of the transition resistance affects significantly the levels of induced interference voltage at the load, connected to the onboard instrument simulator cable.

Transient resistance value increasing reduces the onboard cables shielding efficiency. Thus, while electromagnetic shields designing, it is necessary to account for in shielding efficiency decrease on exposure to thermal and climatic factors during the life cycle of the product.

The aircraft engine start fr om standstill to rated idle in ground conditions can be carried out by electric starting gear fed by either onboard or ground-based power sources. The onboard power sources herewith are accumulators, while voltage can be boosted in ground conditions. Accumulators limit the power of electric start systems. Thus, for starting high-power aircraft engines non-electric systems are used. To increase the power of the ground based engine start systems a high-voltage power supplies can be used. The goal of the article consists in demonstrating the possibility of supercapacitors application to start aircraft engine by electric system. As far as a supercapacitor specific power is greater than that of an accumulator, they can be effectively used for increasing the power rating of aircraft engine start electric systems, wh ere accumulators were previously used. Since the energy accumulated in supercapacitor increases with voltage rise, the supercapacitors can be effectively used in higher voltage systems to increase their power.

The goal of the presented work consists in studying and comparing characteristics and processes of the starting mode of electric motor, fed by power supply containing supercapacitor and without it, based on the results obtained by computer simulation with Electronic Workbench V5.12.

The simulation results confirm the possibility of increasing the aircraft engine start electric system's power by supercapacitor implementation. It means that in many cases it will allow replace aircraft engine start air-compressing and gas turbine systems by electric systems, which possess many advantages and are of great importance for all-electric aircraft.

This work is dedicated to development of experimental test-bench based on magnetic induction rail system. The test-bench allows the ground testing of spacecraft materials and equipment on resistance to micro particles of natural and artificial origin impacts. It will solve the problems related to the costly and inefficient space experiments, and will significantly increase the repeatability, controllability and frequency of impact experiments. In the long-run this accelerator will be an essential part for developing effective protection of the spacecraft from the meteorite hazard, non-existent at the moment.

To solve the problem of low efficiency while converting electrical energy into kinetic energy, which is of great importance for acceleration of small bodies, weighing less than 0.1 g, the authors propose an experimental technique, allowing increasing the efficiency, and, thus, the impactor's maximum speed without increasing the stored energy in storage facilities. The unique feature of the proposed technique consists in increasing the accelerating force acting on the object from external magnetic systems. The authors propose to create a localized external magnetic field directly in the surrounding area of accelerated particles, and then move the magnetization area synchronously with the movement of the accelerated object over the path of the accelerator. This effect is achieved by using multiple-magnetic systems with independent switches and drives, and a single control system. To determine the switching time parameters and parameters of the railguns magnetic systems, the technique of the railgun computation, operating in combination with the multi-loop magnetizing system has been developed.

To test the proposed approach a prototype accelerator was designed and developed. The series of experiments confirming the effectiveness of the proposed method was carried out. Experiments were carried out with particles of various masses, a variety of energy storage levels, as well as for several options for magnetizing systems. The upgraded magnetizing system was 23% more efficient than the classic one, with the same energy storage. The developed accelerator allowed obtain the speed of more than 2100 m/s with a total energy of 11.6 kJ stored in the capacitor bank was reached.

The authors plan to apply the proposed methodology to the main circuit. According to the simulation results, the main circuit multi-step power supply will also contribute to the efficiency increase of rail accelerators.

This work considers the issues of friction (dry, boundary, liquid) effect on the hot die forging process. It reveals the main sources of frictional forces formation.

High temperatures, pressure and permanent renewal of one of the friction deformable metal surfaces being in the plastic state characterize the external friction during the hot die forging. In the course of stamping, as the die fills, the surface area to body volume ratio is increasing. The destruction of oxide films thereupon on the surface of wrought workpieces and the outcome of the non-oxidized metal particles from them occurs. This event facilitates the development of the forces of intermolecular gripping of the wrought workpiece and the tool. The stainless steel, aluminum and titanium alloys are especially prone to sticking to the tool. Thus, their stamping is always carried out with lubrication.

In most cases, the friction at contact surfaces while stamping occurs together with intervening and isolation mediums (oxide scale, oxides, lubricant etc.). Thus, the interaction of lubricants with surface-active substances while stamping becomes of particular importance.

The types of lubricants, their composition and the additives effect on the difficult-to-form alloys of low-plasticity processing are considered.

The mechanism of action of surface-active substances in conditions of stamping and formation of plasticized surface layer with ultra-fine-grained and nano-sized structure was analyzed.

The article analyzed the results of leading Soviet and Russian scientists' studies in the field of nanostructured state forming in the surface layer of the material.

Based of the conducted analysis, we can state that the nano-structuring of the workpieces' surface, including pressure shaping, while applying surface-active substances, leads not only to the obtained semi-finished products' mechanical properties substantial improvement, but also to a significant improvement of their technological properties during the subsequent hot deformation, such as stamping. Thus, the compelling for the production possibility of difficult-to-form materials' super-plasticity deformation under lower temperatures and higher speeds of not only volume nano-structured workpieces, but also the workpieces with nano- structured surface is created.

The study of laminate composite materials, compounds of dissimilar materials and hybrid composite materials for increasing their strength, fatigue strength and reliability is a topical problem for aircraft building.

This work studies the technique for increasing strength of fiberglass with AB-T1 aluminum alloy compound and fatigue strength of hybrid composite material by intermediate layer creation.

To reinforce composite compound intermediate rugged porous layer, obtained by plasma-sprayed method. The paper performed comparative analysis, sel ected materials and modes to such layer formation. Fatigue testing of hybrid composites samples was carried out. Temperature effect on shear strength of a composite compound was studied. The effect of fiberglass molding process (with glue or without it) on the components shear strength.

As a result of the conducted studies we found that:

1. The presence of intermediate layer allows increase shear strength of a AB-T1 + (PN70U30 + EP741) +BK50 + VPS fiberglass composite compound up to 50%, and AB-T1 + (PN70U30 + EP741) + VPS composite up to 90%.

2. Implementation of plasma-sprayed intermediate layer allows increase fatigue strength of fiberglass aluminum alloy compounds up to ≈ 120%.

3. Implementation of plasma-sprayed intermediate layer ensures workability of hybrid composite materials under consideration at temperatures fr om – 60°C to +60°C. The temperature profile  repeats equidistantly the curve of basic technology, but at higher strength values.

The results of the study can be used for new composite materials development and hardening adhesive compounds of dissimilar materials. For example, to develop hybrid composites titanium fiberglass aluminum alloy, and new SIAL variants for fiberglass aircraft propeller blades design, compressor and turbine blades for gas-turbine engines.

Photolithography is one of the main technological processes for obtaining on a special base a certain topology of various electronic components. The most important thing herewith is minimization of all errors in the course of image transfer fr om a photomask to the photoresist layer, and at the developing stage. In this case the most accurate mage transfer is achieved.

The paper is devoted to a topical problem, namely to the photoresist thickness selection specifics to ensure and improve lithography process stability during semiconductor components structures manufacturing.

The paper describes the experimental study of the dependence of radiation energy (E0) dose, necessary to full structures' development in photoresist, from the photoresist thickness (h) on the example of SPR700-1.2 photoresist. The energy dose for the structures' full development in photoresist, determining the quantity of energy affecting the photoresist, required for full photoresist elimination from certain areas, determined by components' structures topologies is one of the basic technological parameters of photoresist.

In the course of the study one area per each of 33 silicon wafers were detected, wh ere the photoresist was completely removed. Radiation energy, at which the exposure of these areas was carried out, is an energy dose, necessary for the full structures' development in photoresist. Thus, the plot of energy dose, necessary for full structure development, versus photoresist thickness was obtained In the course of mathematical calculation, approximation of experimental harmonic dependence was performed and equation of the given curve was obtained.

Rational operating points (thickness) were determined using the plot obtained while experimental curve approximation. These points represent extremums, since with minimum deviation from the rated value, inherent to the considered operating point, the energy dose for full structure development in photoresist would vary insignificantly.

Thus, nine operating points corresponding to a certain photoresist thickness were obtained as the result of the approximated curve analysis.

The result of experimental study of radiation energy dose dependence from photoresist thickness described in this paper consists in obtaining of a number of recommended photoresist thicknesses, which observing can lead to the most accurate image transfer from photomask to photoresist layer, which, in its turn, will improve lithography process stability during semiconductor components structures manufacturing.

The subject of researh in the presented paper are interlayer deffects in composite materials (CM), prevention of the rate of their occurrence by strengthening the compoistes with nano-sized powder. It contributes to safety increase while aircraft operatioin and allows prevent emergency situations duting flights.

The goal of the studies consists in developing methods for manufacturing techmology of samples made of epoxy resin reinforcement of laminated CMs with interlayer defects and nano-dispersed powder, by adding nano-particles to the binder, and obtain maximum degrees of CM filling with nano-sized powder.

In the process of performing this work the experimental samples were produced, and the series of tests were conducted. It is noted that occurrence of interlayer defects contributes to lifespan reductiono of a product made of composite materials. The analysis was performed, thereby, on determining the character of a delamination type defects growth.

The properties such as mechanical characteristics anisotropy and the possibility of hidden defects presence in the form of material discontinuity over the separation surface are intrinsic specific properties of composite materials. The paper presents the experimental results of the study how degree of filling of the ED-22 resin by nano-sized silicon dioxide powder (“Taroksil” T-20) of various concentration affects the mechanical properties of a heterogene material. A brief description of production process technology of samples, made of epoxy resin and nano-disperced powder, is presented. The above said studies are used for solving the problem of interlayer defects hardening in laminated composiste materials, which occurrence is a consequence of the aircraft parts production technology imperfection and effect of operational loads of aircraft, by adding the nanoparticles to the binder. The optimal degrees of CMs filling by nano-sized dioxide silicon powder in dependence of mass concentrarion were found. The testing results of the samples made of CM with embedded interlayer defects with adding nano-dispersed additives with various volume concentration from 0.1% to 0.5% to the binder are presented.

These tests tresults' data would be offered for implementation by the enterprises of Holding JSC “Helicopters of Russia”. The work is prospective for further consideration and implementation in the future research activities.

The scope of the article is production monitoring system developing at the machine-building enterprise in conditions of single-part and short-run production, allowing data provision of production progress, contribution to managerial decisions effectiveness increase and production lead-time reduction.

This goal implies organization of a production processes information system at the enterprise that accumulates initial order information, such as engineering and design documentation, route technological processes, operational labor standards. This data allows calculate a period of execution of works (with account for product structure) and draw an activity network as the basis for the day-to-day production planning. Actual production data is fixed in the strategic points relating to such production stages as resource supply, mechanical processing, finished items transfer to the picking store and to further assembly process. Analytical comparison of the initial and actual data serves as the basis for the management decision-making concerning de-bottlenecking and can be used in production scheduling.

It is supposed to use two groups of Key Performance Indicators (KPI) of production activity that characterize:

– the conformity of production progress with planned, estimated and directive periods of execution of entire work as well as intermediate production stages;

– production volume expressed in terms of production lots quantity or standard hours of work that makes it possible to estimate the current and coming labor content as well as production continuousness.

The key feature of the procedure is that all the data necessary for the monitoring is formed automatically with execution of standard working functions of a person responsible for their execution, disposing of the difficulty to obtain additional manpower resources.

This system serves as the basis for the production processes operative monitoring adapted to the single-part and short-run environment. It allows:

– systematize and analyze the real-time data on the basis of measurement of the strategic points that specify the execution of order;

– control the process on every level, from foreman to general manager;

– estimate the planned machine utilization.

This technology of production data organization is implemented at the DB “Armatura”-branch of FSUE “Khrunichev SRPSC”; the day-to-day production planning and monitoring algorithms are being tested. As a result of the partial implementation of the procedure there is a tendency to reducing of the throughput time. The suggested technology can be used at manufacturing enterprises with a high level of experimental development, such as engineering departments with a pilot plant.

Nowadays developing enterprises undergo hardship to find sources for innovative projects financing. Investors do not want to invest in risky programs, while enterprises are no able to bear the financial risks themselves. A distinctive feature of the high-tech defense developments production financing is preferential financing at the expense of budgetary funds. However, that does not absolve enterprises from financial risks due to insufficient financing or failure of terms. In this case, the enterprise is better to involve extra-budgetary financing. It would help: to exclude cases of forced attraction of own funds; to increase profitability and avoid loss of development; to reduce the probability of disruption in terms and penalties; to release a part of the developer's funds for own technical re-equipment and development; to reduce the impact of design, technological, financial and economic risks.

One of the distinguishing features of the development of high technology products is the presence of the objects of exclusive rights. These objects have a certain value, depending on the area of their application, utility, degree of elaboration and novelty. Additional funds for innovative projects implementation may be attracted by the sale of options for these objects to interested parties.

For the purposes of the developing company, it is necessary to sell such a number of options and at such a price that it would be able to ensure a risk-free return (i = 25%). For the company acquiring the option, the price is defined as the present value of future benefits taking into account the likelihood of a favorable outcome.

Including the price of sold options in the financial model of the project, you can achieve its payback.

In modern conditions of the Russian economy and restructuring of certain industries, the significance of enterprises’ management increases.

It is stipulated by the disruption of a great number of economic ties, active import substitution of products manufactured by certain industries, as well as the necessity for significant breakthrough in the field of military industrial complex (MIC).

Instability of external environment with high risks level is aggravated with account for the challenges facing the Russian industry in conditions of scale sanctions.

In this connection, the role of small and medium-scale enterprises, which significance in the modern world is steadily growing in both developed and developing countries, is increasing.

Methodology for assessing and monitoring industrial enterprises' adaptation level to the crisis conditions is analyzed in real conditions.

Monitoring system of adaptation level of separate enterprises to crisis conditions is an important element of anti-crisis policy, since it should contribute to the selection of the strategically important business-partners selection, subcontractors on production activities, etc. The enterprises with high-level adaptation to the crisis conditions should be primarily included in industry plans while preparing and implementation of the products critical for the industry.

The relevance of small and medium-size enterprises adaptation process is associated with the fact that their ordering parties are large enterprises. So operations of “business for business” (b2b) type are implemented. Thus, small and medium-sized industrial enterprises in most cases are not oriented on individual consumers, but on a big business, which establishes its own rules of conduct on the market.

The area for the study is steady development mechanism of economy of industries, complexes and enterprises.

The object of the work is the Russian small and medium-sized industrial enterprises overcoming the crisis phenomena in the economy.

The subject of this work is methodical and practical approaches to the Russia's industrial enterprises adaptation to negative conditions in the economy, including the assessment of the adaptation level of the enterprises.

The aim of this work consists in developing a methodology of Russia's small and medium-sized industrial enterprises adaptation to the conditions of crisis. The methodological basis of this work is a systematic approach to the small and medium-size enterprises adaptation assessment to crisis conditions.

Practical significance is determined by a comprehensive quantitative approach to the assessment of industrial enterprises' adaptation level to the crisis phenomena in the economy.

In theoretical terms, the overall conclusions on the adaptability of economic systems can be formulated:

1. High level of adaptive properties of any system means that significant changes in the external environment cause insignificant reaction of the system.

2. In the framework of market relations, the more so in crisis, the important characteristics of the external environment are mobility and uncertainty. Thus, the adaptability acts as a fundamental property of such dynamic systems as a subject of small and medium-sized industrial enterprises.

3. Adaptability allows maintain an optimal level of internal processes flow in the system, while the system itself acquires stability and ability to survive in the existing environment.

Aerospace hydraulic systems generate pressure and flow-rate oscillations in the course of their operation, which in its turn leads to vibrations and noise level increase.

Thus, the problems of the study can be formulated as follows:

1. Development of a model of hydrodynamic processes in gear-type pump, accounting for dynamic processes in a locked volume, two-phase nature and pressure oscillations of the working substance.

2. Balancing grooves profile in gear-type pump front foot bearing design, allowing working substance overpressure in the locked volume.

3. Determine experimentally the effectiveness of design procedures on the improvement of gear-type pump dynamic characteristics.

The authors realized numeric model with allowance for the two-phase nature of the flow and the pump's design features using programming language Delphi. Computations allowed obtain the fuel consumption patterns at the input and output of the pump, as well as cavitation phenomenon in the locked volume. Based on computation results, a technique for balancing grooves in front foot bearing was developed.

The effectiveness of such changes in construction was demonstrated experimentally at the Institute of Machine Acoustics. Using scada system LMS Mobile the authors fixated reduction of vibration, pressure oscillations and noise for NMSh-5-25-4 pump.

Thus, all the planned tasks of the research were fulfilled.

The results of this work were implemented at Wroclaw technical university (Wroclaw, Poland) and Institute of Machine Acoustics (Samara, Russia).

Rockets MMP-06 and MMP-06M are nowadays are the most popular rockets in Russia, while the “Dart” system is the most popular rocket in the USA.

The MMP-06M reentry vehicle maximum flight altitude is 60-80 kilometers.

Since 1988 the rocket is employed for the wind velocity and temperature measurement in upper atmosphere.

The rocket is equipped with the engine with a steel body as a thruster, which results in low thrusted efficiency for modern technology state-of-the-art (the empty engine with stabilizers to fuel weight ratio equals 0.56).

The altitude of probing is relatively low, and equals 80 km, and probing at the flight upward trajectory is impossible.

To avoid the above-mentioned drawbacks, the SC Instrumentation Design Bureau under contract with the Ministry of Environmental Monitoring and Research Developed “Mera” meteorological rocket with probing altitude greater than 100 km.

To serve as a thruster the surface-to-air bi-caliber missile engine, using solid propellant (with density impulse of 240 kgFs/kg) was developed and finished-off, with fiberglass body and the empty engine with stabilizers to fuel weight ratio of 0.3. It allows significantly reduce initial weight of the rocket and its size.

The meteorological rocket “Mera” was designed based on the above said surface-to-air bi-caliber missile engine, and MMP-06M “Dart” as reentry vehicle cruise component.

To provide requirements fulfillment (achieving altitudes over 100 km) meteorological rocket “Mera” has two-stage structure with passive cruise component and equipped with a booster.

Measuring and servicing equipment is allocated in the cruise component in the form of a container. The cruise component is equipped with parachute in a separate container.

On the assumption of stiffness conditions and required temperature the body of the cruise component is protected by combined coating.

To ensure radio signal of the equipment propagation the cruise component is equipped with radio-transparent insertion.

To ensure aerodynamic stability the cruise component is equipped with asymmetrical consoles.

The paper presents aerodynamic, weight, inertial and ballistic characteristics, impact zones and separated engine trajectories, as well as cruise component impact zones.

Hall thrusters are widely used for satellite orbit correction and marching operations for altitude change. At the same time the accelerators designed according to similar schemes acquired wide spread occurrence in vacuum ion plasma technologies as ion-cleaning and nano-scale surface treatment systems.

In a first approximation, in the design of such devices it is assumed that the magnetic field does not affect the ions movement in the accelerating channel. Actually, the ions deflected slightly in azimuthal direction under magnetic field impact, whereby the beam acquires the shape of one-sheet hyperboloid. With the thrusters, it might lead to the plume spread, derating and angular momentum occurrence. This leads to significant divergence of the ion beam in the technological accelerators operating on relatively lightweight argon. For surface cleaning before coating deposition such

divergence of circular beam is acceptable, since maximum processing area is required. However, for dimensional ion beam processing narrow ion beams with Gauss ion current density distribution are required. At the same time, effect of the ion azimuthal deviation does not allow focusing the ion beam of the Hall accelerator only by coning the walls of the acceleration channel.

In this paper, additional magnetic pole was installed for focusing ion beam into a spot with Gauss ion current density distribution along radius at the outlet of the cone acceleration channel of the ion source. This magnet pole produced the magnetic field which vector is opposed to magnetic field vector in the channel. Ion beam in the additional magnetic pole area turns in azimuthal direction, opposite to its turn in the acceleration chamber. As a result, the beam is coned and focused at a specified distance into a spot with maximum ion current density concentrated in the center.

The paper formulates the criterion of optimum ion beam focusing in accelerator with anode layer. The ion current density distribution along the radius of the focused ion beam was measured with the accelerator experimental sample. It was shown that the installation of additional magnetic pole allows focusing the ion beam completely.

The obtained results can be used in the design of ion sources for punctual ion-beam machining of the details for optical and electronic industry.

The paper analyzes the concept of distributed power plant (DPP) for prospective long haul passenger aircraft. This DPP is intended to provide deeper integration of a power plant and an aircraft, as well as increase its fuel efficiency.

Possible variants of drive realization for external fan modules, as a constituent part of the distributed DPP, are presented. The necessity of considering the gas drive, realized by introducing an additional transient duct between the turbine of dual-flow turbojet engine and the turbine of the external fan.

The paper presents the preliminary analysis of DPP with gas drive of an external fan module developing possibility in the simplest for realization version incorporating a single external module.

The authors developed the technique for numerical study, carried out the evaluation of the specified DPP parameters under various values of total pressure losses in the transient duct to the external fan module and performed preliminary evaluation of the distributed power plant weight.

Further development of the considered distributed power-plant concept the additional gas heating in the transient duct while the take-off mode is offered. Additional calculations of new type engines are carried out, and estimation of new distributed power-plant structure parameters improvement possibilities is made.

In conclusion, comparison of the considered distributed power plant structures basic parameters at various degree of total pressure losses in the transient gas duct to the turbine of the external fan module is presented. The conclusion is drawn on the necessity to assign transient ducts and intermediate heating systems technologies to critical category.

As usual, thermodynamics and statistical mechanics deal with the problems, which suppose a system to be in equilibrium. Thus, the implemented mathematical tools could be rather conditional in cases of the systems with irreversible chemical reactions, as well as gas flows with thermodynamic non-equilibrium. Depending on how the system differs fr om equilibrium state, the great majority of practical solutions of combustion problems referenced in classical literature this condition is observed. In contrary cases, i. e. significant non-equilibrium in combustion in high-speed flows problems, and detonation in particular, the variation fr om stationary proliferation of chemical reaction fronts is unreasonably neglected. The traditional combustion in gas flows problem statement is unidimensional and based on consumption, momentum and energy conservation laws. Effects of viscosity force and thermal conduction are accounted for herewith. The basic difference of idealization consists in supposition of consistency and averaging of thermal capacitance value under constant pressure and volume. However, these values are dependent from chemical components composition and temperature in particular. Viscosity factor and other factors, characterizing transition are also functions of gas mixtures composition and temperature. As a consequence, gas constant and constitutive equation differ significantly from the idealized form. For complete analytical description of combustion gas dynamics, accounting for mutual diffusion of chemical components, regularities of components vanishing and occurring of new ones, as well as evaluation of total heat emission due to the completion of chemical reactions are required.

Systematic numerical studies of homogeneous and heterogeneous chemically non-equilibrium gas flows in aircraft power station nozzles are already conducted for many years. Various authors obtained results for combustion products of a number of fuels employed in aviation and rocketry. However, calculations of such flows do not satisfy modern practical requirements in all respects. Their main disadvantage consists in orientation on strictly defined set of substances and chemical reactions. To other shortcomings are neglecting the small concentrations of the reacting components, which compels to coarsen recombination mechanism. The variety of propulsion installations designs predetermines the presence of various units with non-equilibrium combustion in the area of lean and reach mixtures, such as gas generators; liquid rocket engines combustion chambers with complex mixture-formation systems; a number of pressurization systems and gas passages with gas flows; combustion chambers and afterburners of air-jet engines.

To a certain extent, determination of specific impulse losses in rocket solid engines due to chemical non-equilibrium with allowance for its effect on formation of Al₂O₃ and ALN condensed particles presents practical interest. The nowadays reality is the study of combustion detonation mode, wherein flows idealization is unjustified due to high conversion rates, and chemical reactions are principally non-equilibrium. The paper presents gas flows with non-equilibrium chemical reactions modeling in the form of conservation equations: uniformity of energy and impulses, wh ere impulses are presented as a product of gas mixtures density scalar and their velocity vector. As a result, in addition to the equation in Navier-Stokes form the authors obtained one more member, accounting for relaxation processes in thermodynamic system. Based on carried out analysis of the law of mass action the authors obtained interrelation between Gibbs thermodynamic potential with the equation member, accounting for non-equilibrium in gas flows with specified content in the form of normalized function. Based on it, the authors offer an engineering design procedure of a rocket engine specific thrust losses (aircraft power plant) caused by chemical non-equilibrium. The values of combustion products equilibrium and frozen compositions for the specified fuels are used for computation of adiabatic coefficients for lim it cases and normalized function. The paper presents graphs illustrating the computations for a wide spectrum of combustion products compositions. The examples of computation results of specific impulse for various cross sections of rocket engines nozzles.

The engineering method for calculation of the thrust specific impulse losses occurring due to chemical non-equilibrium allows estimate adequately their contribution to the common share of losses.

Improvement of modern GTE and power plants directly related to improvement of the combustion chamber. However, combustion chamber is one of the most problematic parts in terms of the design and finishing-out. To solve these problems the authors developed the technique for performing common computations of the compressor and combustion chamber together. In the framework of this work this method was used studying the effect of flow unevenness, occurring behind the compressor blades, and on combustion chamber workflow.

The method has been further developed in the way of implementation of common mesh model for the compressor, the combustor, and working out the boundary conditions setting principles. Geometrical model consists of four different geometrical volumes: guide vanes of the penultimate stage of high-pressure compressor, the impeller and guide vanes of the last stage and the flow path of combustion chamber. The sector of compressor and combustor was used instead of full-sized model to reduce calculation time. The sector angle kept constant for compressor and combustor.

Three-dimensional modeling software package Ansys Fluent was used for simulation of common operation of compressor and combustion chamber, since the combustion processes simulation was tested and verified for this package. Mathematical model and boundary conditions were set after mesh generation. Mathematical model included different calculative models, which were necessary for the combustion simulation. Boundary conditions were specified by temperature and pressure of the flow at the inlet and of the fuel. The flow blows the guide vanes at a certain angle. Hence, the direction vectors were set in cylindrical coordinates. The simulation was carried out in non-stationary arrangement. Thus, the certain time step and number of time steps, which are necessary for convergence, were set. The simulations were carried out for three engine operation modes (nominal, 0.7 of nominal and 0.5 of nominal regimes) with and without compressor. The least effect of the compressor detected at the the engine nominal mode, and the the largest was detected at 0.5 of the nominal. The obtained results were compared with the results from simulation without compressor.

Simulations revealed that that blade wakes extend up to the flame tube head. These wakes change the flame tongue, pressure field, temperature and velocity in the recirculation-mixing zone. It can affect combustion efficiency, ecological performance and temperature field at the combustor outlet. Thus, the simulations, which accounted for combustion chamber and compressor, more fully represent the characteristics of the working process of the combustion chamber and increase the efficiency of new products design.

Combustion of fossil fuel is accompanied by a number of toxic agents' formation. Nitrogen oxide and carbon monoxide are the most ecologically destructive, for they hurtfully affect humans and the environment. For these reasons the paper solves the topical problem on creating a diffusion combustion chamber for a converted aircraft gas turbine.

For the purpose of efficient aircraft engine combustion chamber conversion from fluid to gaseous fuel, the author proposes the combustion chamber design and complex approach, including of engineering and design studies and experimental studies.

The experimental method includes three stages. At the first stage, the butners' outlet parameters are defined. For this purpose, a workbench for determining a burner throughput capacity and obtaining concentration pattern of the air-fuel mixture in swirling jet burner outlet. CO₂ was used as a gas fed to the fuel ducts, instead of methane. Concentrations distribution over the sections after the burner presents the pattern, allowing trace the CO₂ concentration level variation dynamics in whole area of measurements and in each point of the swirling jet. It allows evaluate the quality of air-fuel mixture preparation. The burner throughput capacity was evaluated at various pressure differences. Based on the performed work, selection of the burner geometry for implementation in the compustion chamber was performed.

While implementation of the flame tube head with a large number of atomizers, fuel distribution uniformity ensuring is of especial importance. It provides stable combustion process and mixture homogeneity at the combustion zone inlet. To determine the flame tube head flowrate characteristics, an installation with compressed air delivered to fuel ducts was implemented. Evaluation of air throughput deviation from its average value was carried out. It allowed working out the flame tube head from fuel feed ducts dimensions' optimization viewpoint.

The next stage consists in working with a full size combustion chamber. This stage includes two trends. The first one is the pressure loss determination in the combustion chamber, while the second one is determination of the non-uniformity of the outlet temperature field. Selection of combustor can degree of opening and air distribution along its length to provide optimal pressure losses and temperature field.

At the final stage the combustion chamber as a part of the engine functioning test was carried out. The engine throttle performance characterization and measuring the exhaust emissions of the engine was performed.

In accordance with the results of the studies, conclusions were made that the realized complex approach to toxic agents emission reduction allowed design the combustion chamber reducing nitrogen oxide emission by 40% and carbon oxides by 20% compared to a stock combustion chamber.

With implementation of existing methodological support of regular automated verification systems (AVS) the state and performance of a short-range missile of air-to-air class (AAM SRM) control system sensors are unobservable. Thus, while regular AVS typical control algorithms realization technical state of control system sensors, such as linear accelerometers (LA), or angular accelerometers (AA) can be estimated through indirect parameters, without their basic parameters determination (transfer factor, etc.). This could significantly reduce methodological fidelity of guidance system control.

To solve the above said problem the authors offer implementation of functional control (FC) method. This method can be realized based on software/hardware complex (SHC).

The paper suggests scientific basics of functional control. They are stipulated by implementation of harmonic balance of automated control theory. The FC structure, organized by duplication method, was used to realize AAM SRM guidance system FC control sensors.

To minimize the control structure dimensionality at the inputs a single primary impact  on the missile guidance system is applied using harmonic oscillation workbench (HOW). To close FC links one should be aware of HOW functioning as a master side of SHC.

HOW is the main preset part of SHC, generating a single primary stimulating effect  on the missile during FC of its guidance system sensors. To close FC system it is necessary to set correct stimulating action on an FC object. It is necessary herewith to eliminate FC resonant mode, and ensure FC main sensors functioning in linear range, i. e. exclude: guidance system signals overload limiting for LAs; missile body spin velocity limiting for AA detection unit, as well as angular target tracking rate and locating angle limiting for target-seeking head (TSH).

To ensure harmonic oscillation “comfort mode” for the missile guidance system, selection and adjustment of HOSs design values is carried out. For this purpose, the developed SHC FC master side model is used. In addition, the developed model is used for characterization of secondary stimulating effects on HOW, LA and AA detection units and determination of the signals of their reactions.

The process of HOW operation can be represented by a certain model in Laplace operator form. This model includes oscillating and measuring loops. The oscillating circuit dynamic model represents an oscillating link with time constant and damping factor, as well as nonlinearity of saturation type with known parameters, stipulated by HOW design specifics.

Measuring loop includes axial power transmission (PT) and inertia-free angular sensor. PT is free of reduction elements, and its gain K_PT = 1.

The conducted experiments on a certain HOW embodiment confirmed the adequacy and performance capacity the developed models of SHC FC master side, as well as correctness of the HOW design values, allowing eliminate AAM SRM guidance system's signals limiting and their termination to stop.

The authors conduct studies of gliding unmanned flying vehicle (UAV) flight dynamics at low speed and launch altitude. In the case under consideration the UAV flight dynamics significantly depend on initial flight speed and initial flight altitude which determine the total UAV energy and, as consequence, UAV's dynamic capacity while moving along the trajectory.

The paper considers the following two problems:

1. Maximum flying range provision under initial UAV motional energy limitations.

2. The UAV stability and maneuverability provision at all flight stages.

We assume the UAV is equipped with a certain booster engine with fixed total impulse, which can be realized by various thrust variation functions in the course of UAV movement.

Much attention was paid to the study of launching conditions and thrust behavior at the initial trajectory portion impact on the flight range under gross thrust impulse limitation, as well as studying of various possible technological deviations of thrust vector direction from UAV axis of roll impact on movement stability and UAV launching safety. The last problem was considered in the form of the following two problems:

• ensuring such UAV angular stability at the initial passive trajectory segment, which would guarantee UAV angular orientation, eliminating the possibility of UAV collision with the carrier by the time of its engine firing;

• ensuring the possibility of disturbances parrying, which occur during engine operation at the active trajectory segment and stipulated by technological errors of its mounting on the UAV. The results of the study revealed that the last factor could affect negatively as well on the UAV total flight range.

We assume that the UAV is launched in undisturbed air conditions so that at the starting moment it is not subjected to the additional aerodynamic impact, while the carrier is moving at constant altitude with constant speed. The authors developed a mathematical model of UAV spatial motion all over the flight. The control system accounts for pitch angle and angular velocity deviations. Solid fuel accelerator with fixed thrust impulse value, variable thrust value and operating time is considered as a boost engine. A time of engine ignition was computed. Movement parameters at the initial trajectory segment, booster thrust variation functions impact on the flight range and booster thrust misalignment impact on the UAV movement parameters and stabilization were evaluated. Extreme (guaranteed) values of solid fuel booster thrust misalignment caused by technological errors while booster manufacturing and mounting on the UAV ensuring the UAV flight safety at two stages – controlled flight without thrust and controlled flight with operating booster were obtained.

One of the serious problems of flight simulation with flight simulators consists in reproduction of accelerations experienced by a pilot while in flight, which play an important role in piloting. The paper considers this problem in the context of aircraft in-flight refueling. The goal of the study is quality estimation of cabin movability over various degrees of freedom effect on piloting, pilots nature of action and his judgement on simulated accelerations degree of adequacy to real flying conditions.

Experiments were conducted with TsAGI PSPK-102 flight simulator containing cabin with six degree- of-freedom mobility, collimated visualization system, instrument display, side-stick control with electrical loading system, and thrust control levers. The authors developed the in-flight refueling task simulation technique using flight simulator with movable cabin. The problem of cabin mobility system control algorithms optimization was fulfilled for the considered task.

The pilot's task consisted in performing closing-in with the refueling tanker and carry out the refueling cone in the course of the flight. Experiments were conducted with participation of an Honored military pilot, who had wide practical experience of refueling tasks in real flight conditions.

Experimental data on the accelerations effect on unbiased indicators of the cone tracking accuracy, pilots actions characteristics and aircraft movement parameters were obtained.

The study demonstrates that reproduction of accelerations affecting a pilot significantly increases the adequacy of in-flight refueling problem simulation to a real flight. According to the pilot's, opinion axial accelerations exert the strongest effect on refueling task.

Nevertheless, reproduction of vertical and lateral accelerations in the course of flight simulation plays an important role as well. The obtained objective data and the pilot's opinion accord well with overloads and angular accelerations over various degrees of freedom significance analysis performed based on earlier developed theoretical approach to the accelerations impact on piloting.

Lithographic processing is one of the key operations of technological processes while semiconductor devices' and integrated circuits manufacturing. Its parameters effect strongly the precision of the devices structure creation, and, as a consequence, its output characteristics. Multi beam lithography is implemented in particular. Its technological equipment uses silicon electronic lenses for electron beam control, which electronic and optical parameters affect the accuracy of the manufactured product structure and, as consequence, their output characteristics.

The paper tackles the topical problem of ensuring the specified electro-optical parameters of electrostatic lens (including geometric sectional shape of electron beam) during its production, storage, and transportation, as well as repeatability of these parameters from batch to batch.

The research object of this project is electro-static lens representing silicon plate with a plenty of holes of circular shape. The lens under study is used in technological equipment for multi-beam e-lithography for a powerful beam splitting into a multitude of beams.

The electro-static lens parameters degradation causes in length of time identification, and their elimination technique development are the main tasks of this studies.

In the course of the study, a number of operations and factors that could affect the electro-optical lens parameters was revealed. According to the results of expert evaluation of electronic lens manufacturing technological process, these factors are oxidation and chemical cleaning operations.

The results of various technological operations and factors effect on electro-optical lens parameters variation were presented. While this research a series of experiments was conducted, which considered variation of electro-optical lens parameters in length of time.

The obtained results of the studies allowed revealing possible reasons of electro-static lens parameters degradation in length of time, and developing technological recommendations to prevent this degradation.

The plan of future studies is presented.

A voluminous assortment of parts, characterized by higher requirements to accuracy and reliability, is used while aircraft manufacturing. They are fabricated fr om various materials, such as steel, aluminum, plastics and composites. Special attention is payed to developing new methods of the parts manufacturing and improvement of conventional technologies.

The majority of the parts is produced by pattern cutting of various materials of 0.5 to 200 mm thickness, followed by their machining or without it. It is interlinked with the development of CNC metalworking Machinery Park, where forged pieces or form workpieces are used in increasing frequency. The main question consists in productivity and quality of cutting blanks. There is a great variety of pattern cutting methods, distinguished by productivity and principles, with peculiar advantages and disadvantages. The authors consider the hydro-abrasive cutting, which is the newest and prospective metal cutting methods.

Hydro-abrasive cutting is the most up-to-date and efficient method for obtaining either blanks or parts from plate aviation materials. The cutting process is carried out by the thin water jet with abrasive grains mixture, emitted at high (supersonic) speed under high pressure up to 6000 bar. Garnet sand with 7.5-8 hardness is used as an abrasive material. The process represents erosion destruction under impact of working jet, wh ere the abrasive cuts the chips microlayers, while water takes them away from the cutting zone. The main advantages of hydro-abrasive cutting are high productivity ensured for high cutting speed (steel up to 300 mm), the absence of residual strains at the cut edge, the possibility of cutting practically any metal and non-metal as well as the ability of cutting figured profile and irregular shape parts.

Nowadays the process of hydro-abrasive cutting is poorly studied. Theoretical dependencies accounting for all technological parameters effects for the cut ruggedness and corrugation determination, and dependencies reflecting the value of cutting jet lagging.

The quality of hydro-abrasive cutting depends on the feed rate, the thickness and type of cutting material. It was found, that feed increase reduces the quality of cutting, increases ruggedness, and the area of smooth

cutting reduces, while the corrugation and obliquity of the cut increases. Deffects caused by jet lag cutting, such as formation of a burr on the sharp outer corners, forming holes in the inner corners, overcut and undercut at the beginning of the cut are also found.

The goal of this study was to explore the effect hydro-abrasive cutting modes, namely the feed effect on the cut roughness.

After a row of experiments the samples made of three different materials with 30 mm thickness, namely, steel 30HGSA, aluminum D16, multi-layer polymer composite such as titanium-fiberglass were obtained.

When cutting the feed was changed stepwise from 5 mm/min to 120 mm/min for a sample of steel, to 200 mm/min for samples of D16, and to 160 mm/min for a sample of the composite. The ruggedness of these samples was measured at the specific areas of the cutting section.

Analysis of ruggedness dynamics allowed suggest a mathematical model of cutting surface ruggedness profile forming. The ruggedness is formed by free abrasives, which remove repeatedly the micro-chip layers. The mathematical model is proved by experimental data, as indicated by a graph of the cutting ruggedness dependence from the cutting head feed.

The experimental data and theoretical curves allow predict the cut quality of the hydro-abrasive cutting. Based on this data, the possibility arises to select the most optimal hydro-abrasive cutting mode cutting, or a certain type of defects elimination. The cutting rate optimization is possible by slowing down the feed in areas of defects formation, or ruggedness unevenness.

Irregular shape items manufacturing from polymeric composite materials (PCM) requires the tool set, which geometry duplicates geometry of the item. The material is spread on the shape-generating tool set, and then its polymerization is carried out at the predetermined pressure and temperature that can achieve up to 200°C. In this respect, the most complicated problem while forming precision items from PCM consists in temperature deformation occurrence while polymerization process.

For years, metal hybrid tool sets have prevailed in high-precision composite parts manufacturing. A hybrid tool set has invar (nickel alloy with CLTE close to zero) shaping plate and a support structure made of some other metal with sufficient thermal conductivity. The tool set of such kind involves shape-generating plate attachment to the support structure means, which ensure the possibility of their free thermal extension. The drawback of metal tool sets consists in their high cost, low material utilisation ratio and long manufacturing cycle.

The next step in tool sets for high-precision items made of PCM evolution was creation of composite shape-generating tool sets. Fiberglass and carbon reinforced plastics are implemented for such tool set manufacturing. Its surface can be coated with ceramic or gel coat layer of precise thickness, providing minimum roughness, maintainability, and increasing the items takeoffs. Composite tool sets does not have disadvantages of their metal counterparts, though several design problems are still stay unsolved.

This paper proposes a carbon composite tool set design for satellite antenna reflector producing. The main requirements to this tool set are precision and stability of the shaping surface. Design solutions are validated by thermal and static mechanical analyses based on finite elements method. In addition, the paper presents the results of autoclave operation simulation, which allows analysing the tool set optimal positioning inside the autoclave to provide uniform heating.

Aircraft engines manufacturers face the following challenge: on the one hand to ensuring the product configuration management, and, on the other hand, the necessity of effectively employing financial, material and human resources throughout the entire aircraft engine life cycle. This problem can be fixed due to information support of aircraft aviation engines' life cycle based on software solutions such as Product Lifecycle Management (PLM) and Enterprise Resource Planning (ERP).

The aircraft engine building branch intensively employs PLM solutions developed by Siemens PLM Software at the stage of design documentation development. aircraft engines design is carried out with allowance for the methodology of 3D electronic model design (EMD) of a product with NX system under control of PLM-system TEAMCENTER.

PLM solutions are not used at the stage of the technological preparation for assembly production. The technological preparation process is oriented on implementation of paper design and technological documentation. The existing process does not link the stages of design and technological documentation development. At the stage of technological preparation of production process, the electronic structure of the product is practically never used. The technological preparation and configuration management systems depend largely on the human factor. The production planning system is not working effectively due to the absence of connection with technological regulations and assembly production process.

The product assembly efficiency can be improved by creating a unified information environment for developing design and technological documentation.

The TEAMCENTER PLM system implementation for technological preparation of aircraft engine assembly production will allow develop a unified information environment for developing the design and technological documentation. It will enable also the product's configuration management problem fixing and reducing time and costs associated with aircraft engines.

A key feature of the new business process is the TEAMCENTER system implementation at all stages of production technological process preparation, and the products configuration and assembly are carried out according to electronic technological structure and technological process using 3D visualization and step-by-step account of assembly process. Technological structure of the product will allow fixing the problem of production configuration management throughout assembling process. The technological structure data should be transferred to EPR system of production planning.

The equipment and assembly must be carried out via the MBOM and technological process using the three-dimensional visualization and operational accounting of assembly production process. The operational account will allow monitoring the production progress and providing feedback to the production planning system. The step-by-step account will ensure documenting of the product configuration requirements carrying out and forming actual product configuration.

Introduction of the TEAMCENTER PLM system while preparing the assembly production will allow solving the configuration management problem.

Formation of actual configuration will provide a solution to the product configuration control problem, such as documentation, identification, and traceability of the requirements compliance status to products at all product assembly stages.

Implementation of the 3D product models facilitates understanding of new products design, allows exclude drawing working documentation to automate the technical documentation development process and, as a result, reduce the time of assembly and manufacturing costs.

Developing a unified information environment for design and technological documentation preparation via the PLM TEAMCENTER system will provide the market launch of new products with the specified characteristics in the shortest possible time and at the lowest cost.

One of the promising trend of modern transportation systems development is transition to electric propulsion. This is topical for aircraft industry too. However, to solve this problem it is necessary to design electric motors with high power density over 20 kW/kg. To achieve such figures of the specific power is possible only using cryogenic cooling, and modern superconducting materials.

Design of the electric motors with superconducting inductor and armature windings, will allow obtain maximum benefits in terms of weight and size. This relates to the possibility of increasing the magnetic induction value in the motor air gap, as well as with the stator linear load increase.

Design a fully superconducting electric motors is complicated by the absence of any universal computation methods, as well as a number of design features and the critical parameters of high temperature superconducting tapes nonlinearity. All this requires the development of new computation methods for such kind electric motors.

The paper presents a fully superconducting electric motor with a ring armature winding and the method of determining the its specific power and the results of finite element modeling in three-dimensional formulation.

The obtained analytical expression for the main magnetic flux allows derive an equation for the power density of HTS machines with annular armature winding. It is shown, that this power may exceed the value of 20 kW/kg.

While increasing the aircraft degree of electrification hydraulic drives fed by centralized fluid power systems substitution by off-line electric drives is assumed.

Translational motion power actuators with ball-and-screw gear are widely used nowadays in aircraft flaps, slats and adjustable stabilizers control systems, and operate reliably for a few minutes per flight.

In the absence of strict requirements to the dynamic characteristics of electric actuators, such as high-lift drives, simple electromechanical actuators with controllable electric motors and mechanical gear are already in use.

During the flight of an aircraft, controlled airfoils are exposed to varying loads under the influence of airflows. These loads cause significant mechanical stresses in the electromechanical actuator, leading to accelerated wear of mechanical actuator components. Another problem with the existing electric actuators is their excessive weight and size as well as difficulty to ensure compliance with the stringent operational safety requirements.

Thus, the goal of this research consists in eliminating these deficiencies and improving the energy and operating characteristics of electric actuators. It is necessary herewith to consider the operation of an electric drive either in active mode, when the energy is spent to set the running gear in motion, or in passive mode, when the running gear is fixed in a certain position and exposed to significant mechanical loads caused by aerodynamic forces.

Based on the presented aerodynamic forces calculations, we analyze the designs that solve the stated problems. These designs allow implementing both the passive and the active electric actuator modes.

We propose a design that makes electric actuators more reliable and durable while operating in the passive mode. This is achieved by removing the output arm from the deadlock position to allow a limited range of deflection and by damping vibrations and oscillations caused by aerodynamic forces within that range.

However, oscillations damping by electromechanical dampers is not always efficient, since it may result in weight and size figures increase under high mechanical loads. This problem could be solved by implementing the electric actuator structure with flexible coupling between the ball-and-screw gear and remaining actuator components in the form of modified elastic compensating clutch. This proposed flexible coupling demonstrates small weight and size figures compared to with electromechanical dampers under heavy loads. Thus, such structure can be realized also in spacecraft.

Judging from the above said, the considered electric actuator construction arrangement allows reduce its weight and size figures. The resource increase in electric actuator passive operation mode is achieved by eliminating rigid fixation of the output arm in dead spots and limited oscillations of the output arm in the operating range of a position sensor.

Under real operation conditions of electrical industry products, the degradation variance of their features should be allowed for. The subject of research is various kinds of electrical converters which even slight degradation results in serious technogenic disasters. The paper considers and suggests basic principles for such type of problems solution, and establishes a number of degradation variances regularities.

Characteristic parameters variances analysis allows separate out four most characteristic types of functions for the named regularities description: entire irrational functions or polynomials, fractional rational functions, and functions for processes with description. Evaluation of degradation variances simulation results supposes tabulation of the measured values and selection of such an approximating function which would ensure it the least mean square deviation fr om the tabular dependence. The OLS method ensures the best results for solving the problems of such type.

Analysis the considered functions, describing the degradation process, allows state the following: all functions have the initial value, known for the unit in use fr om the its datasheet. Thus, in the course of degradation variance studies it is expedient to examine only the function degradation variations, instead of the whole function. Initial value of the deviations function equals to zero, and its plot passes through the origin of coordinates. While determining the number of parameters of approximation functions, their number would be one less for deviations function. Thus, the order of normal OLS system reduces.

The residual resource prediction was performed based on solving non-linear equation, wh ere degradation deviation function takes the normative allowable values. While solving the equation, the function arguments lim it value should be defined as an instant of failure. The evaluation of the residual resource was performed based on the instant of failure.

The main task of the article consists in electric power grids basic emergency modes, leading to rolling blackouts, analysis, and high-speed reactive power smooth regulator (RPR) design for the existing domestic reactive power compensation systems (RPCS).

Failures analysis at industrial enterprises and substations revealed that one of the main reasons resulting in avalanche failures and blackouts is reactive power circulation, representing an integral part of complex electromechanical mechanisms functioning. However, the excessive amount of reactive power and circulation leads to complications and serves as the cause of failures.

Various reactive power-compensating systems, such as either static (capacitor installations), or dynamic (synchronous compensators), are widely used in the industry. However, preference is given to the static ones due to low price and durability. Their implementation for the most part pursues only economic benefits, namely energy cost reduction. Nevertheless, with certain updating these installations can be implemented successfully for failures, resulting in blackouts, prevention.

On the example of compensating installation, developed by the authors, they suggest to replace one of the critical elements for the purpose of regulating properties improvement. This element is Reactive Power Regulator of foreign manufacture. It has a number of disadvantages, which presence may result in cascade fault.

The proposed new regulator is a thyristor reactive power regulator consisting of a transformer, network mode sensor, regulator, control block, thyristor switches and a filter.

This paper presents the schematic diagram and computation algorithms for voltage level sensor parameters, network mode sensor parameters and filter parameters. The computation of the unit reliability is presented either.

The paper presents the design and experimental research of the high-speed slotless generator scalable prototype with strip-wound stator core, integrated into auxiliary power unit. The experimental research and computer simulation of the scaled-size prototype in no-load and on-load modes were conducted. They revealed that this generator demonstrates minimum rotor losses and voltage ripples, as well as high specific energy characteristics. The high-speed slotless generator scalable prototype computer model was developed with Ansoft Maxwell software. Experimental data deviation from computer simulation results does not exceed 5%. From the results of scalable prototype computer simulation a full scale computer model of high-speed slotless generator was developed. The main parameters of the high-speed slotless generator were defined and compared with the parameters of the slot-type high-speed generator. The comparison revealed that the slotless variant demonstrated lower losses (by 600 W) with minimal weight and size parameters (not more than 0.2 kg/kW), high efficiency, minimal negative high harmonics effect, absence of the slot ripples, and the simple production technology. Thus, the obtained data shows that the high-speed slotless generator with he strip-wound stator core made of amorphous alloy can be implemented as the generator integrated into the auxiliary gearless power unit. It proves also the possibility of its application in aircraft industry.

The paper presents algorithms for voltage generation at induction motor (IM) windings in vector pulse width modulation (PWM) mode while IM rotation frequency regulating. These algorithms ensure smaller amount of computation, allowing eliminate through currents in the inverter power stage. Vector PWM (VPWM) employs 8 states of inverter switches for IM control. The paper considers the possibility of extra intermediate vector states of the switches, which would not cause through currents occurrence in the inverter and allow include them into IM speed control algorithm. For through current elimination in transition periods between zero vectors and basic and non-zero basic vectors the authors suggest implementation of intermediate switches conditions of the switches, which will be operated on as vectors. Let us consider treat these vector as variables. The authors analyzed the usage of a group of vectors V₀₁, V₀₂, V₀₃, V₀₄, V₀₅, V₀₆ or V₁₀, V₂₀, V₃₀, V₄₀, V₅₀, V₆₀. It allows obtain two most promising algorithm. To control IM output voltage and frequency parameters employing PWM mode, the assemblage of the transitions between acceptable ones, when at least one syllable of a control word would be inverted, will be referred to as vector subset of permitted dangerous bilateral transitions. The simultaneous switching of both switches of a totem pole corresponds to the dangerous transition. In the suggested algorithms, the transitions are in parallel with Karnaugh map sides, which means that they do not cause through currents. Vector PWM voltage and frequency parameters will be considered only in the time interval corresponding to the half of the sector.

Simulation was carried out varying time interval of vector V₀ existence for regulating generated voltage value under invariable T_V4/T_V6 ratio (for s the first sector). The effect of n parameter on the quality of IM drive in VPWM mode. The simulation was performed with MATLAB Simulink. Simulation results are presented.

The paper studies the possibilities of using acoustic emission and gas-dynamic processing (thermo-acoustic processing) methods for initial check of a material (titanium alloy) instead of a traditional method, i.e. optimal heat treatment mode selection.

Implementation of thermos-acoustic processing as an extra treatment of TC6, BT16 and BT23 alloys, demonstrating low mechanical properties in the initial state, ensures grains refining and improvement of property package up to the required level.

Physical features of titanium alloys and specifics of obtaining semi-products on their basis require that a manufacturer should know and allow for these semi-products initial state (mechanical properties, microstructure, etc.) while process design. Thus, the manufacturer should possess the technique allowing promptly estimate and correct mechanical-and-physical properties of the basic material, and in certain cases of a complete spring either.

For the experiment, the authors sel ected the alloys fr om various foundries (mechanical properties, microstructure, etc.).

The presented study area of application is titanium alloys implementation for springs, employed in airspace and other special equipment manufacturing, where the quality of basic material predetermines largely the quality of a final product.

The carried out studies in the area of the basic material quality in spring production allows draw inference on the possibility of a certain initial check modernization, as unattainable part of a component manufacturing process. It is established, that acoustic emission method allows qualitatively estimate the microstructure without labor consuming estimation methods and take a decision on treatment schedule of manufacturing process. ATAP implementation as an extra processing of TC6 alloy, demonstrating low mechanical properties in its basic state, ensures grains refining and improvement of property package up to the required level.

The results of monolayer elastic and thermos-elastic characteristics identification in carbon composite samples, produced by employing of epoxy matrix containing 0.2 wt.% of fullerene soot are presented. The composite samples with reinforcing schemes [0₂/90₄/0₂], [+45₂/-45₄/+45₂], [0₄], [90₄] were fabricated by vacuum shaping. The fullerene soot was preliminary added to a binder and disperse using mechanical and ultrasonic mixing.

The composite monolayer properties were obtained based on the analysis of the results of mechanical tests of the samples with various reinforcing schemes and inverse problem solution. The multilayer properties valuations were obtained, using micro-mechanical, analytical and numerical modeling and solving corresponding averaging problems. Mori-Tanaka averaging method was used for analytical computations for cylindrical embedding problem. Numerical calculations were performed using finite elements method at representative fragments, containing unidirectional fibers. The computations used initial matrix properties values obtained from the experiments, and matrix containing the fullerene soot.

The paper demonstrates that the results of numerical and analytical computations performed to evaluate the unidirectional layer properties are sufficiently close to each other. It follows from these computations that in case of impurities agglomeration, addition of nano-filler should lead in the first place to transverse elastic modulus increase and monolayer shear modulus due to matrix tightening. Pitch module should vary insignificantly since it is defined by filler properties. With the filler addition, the monolayer Poisson ratio practically should not change. These results do not correspond with the experiment, except shear modulus increase. Unlike the predicted monolayer transverse elastic modulus increase, the experiments revealed its decrease. It follows from the experiments that monolayer Poisson ration significantly decreases, which was not predicted by computations. The obtained results demonstrated the matrix embrittlement while implementing the selected nano-modification technique and the necessity of either filler volume fraction decreasing, or changing the technique of its dispersing in the binder.

The authors plan to use identified values of composites' monolayers elastic and thermos-elastic characteristics hereafter to describe the residue stressed-deformed state of carbon composite construction elements to reveal the possibilities of reducing residual stresses and shrinkages in the structures with asymmetric reinforcing schemes, using matrixes containing carbon nanoparticles.

A line of composite materials based on low-density linear polyethylene (LDPE) thermoplastic matrices, polypropylene (PP) and metallic nanoparticles was produced by mixing in polymer melt. The results of dynamic mechanical analysis of PP based composites with metallic nanoparticles, namely the product of Co (II) acrylamide nitrate complex and 2% FeCoAAm co-crystallizatant thermolysis, within the temperature range from −50 °C to +150 °C revealed, that low concentration of nano-filler (1 wt.%) does not lead to noticeable changes in dynamic elastic modulus, nano-composite mechanical losses and loss tangent. Thermooxidative degradation results indicated the increase of thermostability for above said PP-based composites compared to the initial PP at 4 and 8 wt.% of nanoparticles.

The authors obtained nanocomposite materials based on polyolefin matrix and pre-synthesized by chemical co-deposition magnetite nanoparticles such as LDPE-Fe₃O₄ and PP-Fe₃O₄. According to X-ray diffraction analysis, the major component in the system was magnetite nanoparticles with an average size of 15 nm. These results correspond to scanning electron microscopy data. The paper demonstrates that with the increase of nanoparticles content in polymer, and with magnetite high content in particular, the elastic modulus increases, and the tensile strength value decreases. Thermal behavior analysis in the PP-Fe₃O₄ (at 4 wt.%) system indicates that nanocomposite thermo-oxidative degradation reduced compared to the initial polypropylene, and the temperature of maximum degradation start-up increases from 300°C to 385°C.

Composite materials based on LDPE and Al₆₅Cu₂₂Fe₁₃ with alloy (0.1 to 10 wt.%) were produced. The paper demonstrates that the presence of quasi-crystalline alloy as a filler leads to composites strength properties improvement. Unlike LDPE-Fe₃O₄ systems, a tensile strength of LDPE-Al₆₅Cu₂₂Fe₁₃ increases with low filler concentrations.

Protective action of the nanocomposite systems under test in relation to beta-radiation was studied using dose metering method. It was demonstrated that with filler content increase in LDPE-Al₆₅Cu₂₂Fe₁₃ and LDPE-Fe₃O₄ composites beta-radiation flux attenuation occurs. A high correlation between the portion of passing beta-radiation and relative dielectric constant of composite materials based on thermoplastic polymer matrix with metal-filler was observed.

The paper tackles the topical issues of staffing training and forms of factory organization at the aircraft building enterprises based on lean production concept. While production development in Russia and its share increase in the global market, the issue of product company optimal management comes up. The lean production program leads to creation of learning organization with stable, continuously progressing processes, aimed at searching for non-productive losses and their minimization. The lean production becomes the topmost factor of efficiency increasing, competitive stability of an enterprise and reliable technique for all kinds of all kinds of expenditures. In an aggravated competitive struggle at domestic and world markets, the key factor of Russian engine-building companies' success is associated with their flexible response to rapidly changing market demands. This requires development and implementation of a number of measures aimed at improving the efficiency of production and enabling enterprises to enter the global market as providers of competitive aircraft engines.

The main problem at domestic engine-building enterprises consists in production systems modernization. Many companies take the mass production concept as a basis of their production system, which does not meet modern industrial requirements to goods and services production, and does not take the expected effect. Thus, it is necessary to carry out the production system modernization, taking more efficient and productive system as its basis, engaging all management and stuff of the company in this process.

Effective organization of production at aircraft engine-building enterprises is a significant and special component of the competitiveness analysis due to its magnitude to production encompassing and time scale parameters of their implementation. Thus, the organization of competitive aircraft equipment manufacture should allow for all kinds of losses and expenditures, and implement efficient production system, including the great majority of methods, techniques and tools.

The paper tackles the issue of the cost management methodological approaches enhancement.

The main purpose of the paper consists in developing the cost management financial mechanism based on EVA concept for the space-rocket industry enterprises as an instrument of increasing the enterprise financial stability and competitiveness.

System approach to financial management theoretical basics summarizing and analysis composes a methodological base of the research. While conducting the research the author employs the methods of financial analysis and forecasting, economic-mathematical modeling and expert assessment.

The author identified space-rocket industry enterprises' specifics and explored cost management methods existent in economic science. The economic value added concept is a method on which basis the financial mechanism of cost management is developed.

As a result, the indicators of the EVA concept were modified. The adaptation of the basic indicators of the EVA method to the cost structure of the space-rocket industry enterprises and the technique of calculating EVA are specific to the organization of production and the budget process. They allow identify the basic elements of economic value added cost, i. e. purchase of materials (EVA_м), staff salaries (EVA_т), equipment handling (EVA_а) and other expenses (EVA_п), characterizing the efficiency of rocket space technique production process.

Moreover, the optimal cost financing structure was identified. The results of the research were proved on the example of the space-rocket industry enterprises. Finally, we conclude that the most expensive source of funds are borrowed funds, which effective use will consist in material and other costs financing at their expense, but costs of labor and depreciation deductions is preferable to be financed from the own funds.

The paper is devoted to the issues of cargo air transportation market development in Russia, and exploration of ramp cargo airplane An124-100 operation in the market. At present, this plane carries out the major part of transportation of heavy oversize cargo. The plane demonstrates a unique combination of capabilities, as it is one and only air transport for oversize cargo alternative overseas transportation. The plain allows also increase safety and reduce damage probability of cargo, compared to overseas transportation. It is capable of delivering cargo to far-out regions, where there are no auto-road and railways, horizontal loading and unloading capacity through nose and tail ramps, as well as lowering the aircraft floor and unloading without specialized external cargo-handling equipment.

To forecast economic indices of the plane of a specified type simulation model run by Vensim program. The conclusion is drawn that with growth of heavy oversize cargo air transportation demand, the necessity for organizing in Russia the serial production of aircraft equal to An-124-100 increases. Analysis of cargo transportation world market modern tendencies revealed apparent significant potential of the sector of economy in question and its direct interrelation with such factors as the degree of technological development of the country, the State participation in the trading processes and general level of economic development. The branch of group air transportations should be considered as one of aircraft industry strategic orientations totally and enjoy the State support.

The paper considered fatigue resistance testing of helicopter structural elements on the example of helicopter rotor blades samples testing. Endurance testing of aircraft equipment components and structural elements consists in laboratory reproduction of external disturbances corresponding to the standard operating conditions, cyclic loading and functioning. However, these tests do not include studies related to the gradual damages accumulation leading to cracks initiation and propagation and finally to structural damage. In this regard, studying the process of cracks growth while full-scale tests of the samples presents special interest. The paper presents the brief description of blades full-scale tests process with concurrent video shooting. The samples are subjected to static loading, with subsequent additional bending load moment of variable sign. Video records of cracks growth were processed, and data on the crack subcritical growth time was obtained. This information is presented by the diagram, illustrating the crack growth time dependence of the crack growth rate. The paper analyzes measuring and test equipment used while testing for recording values of tensions occurring in the studied samples, due to bending load of variable sign applied to them. Fatigue resistance characteristics were determined, and fatigue graph was plotted. Arithmetic mean and root-mean-square deviation of endurance limit stress are obtained also.

The accidents of two Earth satellites collision when impact velocity of the spacecraft abeam reached 10.5 km/s. This velocity is several times than that required for a crystal lattice inertial explosion of the metal, constituting the spacecraft body. Inertial explosion parameters of metal components, which can occur at the contact point of the collided spacecraft, are studied. The paper demonstrates experimental and computed data on the collision velocity, causing such an explosion, as well as motion speed and explosion vaporous products temperature, reaching 22 000 K. It shows that the time necessary for metal transition from the solid state to luminous atomic-vaporous mixture reaction excitation does not exceed 2 µs, if this transition was caused by mechanical shock. Mass ratio of the exploded metal was determined. All experiments were conducted using lead samples.

Metal preserves its solid state until the metallic binding energy  is enough to preserve its crystal lattice. This energy equals to the sum of a metal heat content from the temperature T = 0 K plus evaporation heat up to the sample sublimation. Acquiring the energy of  the metal ceases to be a condensed media and passes to high temperature vapor condition. Such transition occurs while siderite or nickel meteorite collision with Earth, or spacecraft.

The experiment procedure was as follows. The lead ram tester of a cylindrical form weighted 0.027 kg, had the diameter of 14.5 mm and length of 15.2 mm. Its velocity was v = 1128 + 14 m/s. The lead target was of a parallelepiped shape of 67 × 82 × 15.5 mm and weighted 0.91 kg. The target mass remained after the lead ram tester stroke was 0.68 kg. The rest lead target mass (as well as the ram tester) evaporated.

During the experiment, the velocity of moving elements was determined by images movement on video frames, recorded by Phantom V 16 model 10 video camera. The exposure time per one frame was 1 / 156000 s^-1, and the shooting speed was 25 000 frames per second.

The shock waves pattern in inertial explosion vaporous products of the two lead structures was obtained. The Mach number measured in the open air equals 2.36.

The paper considers perspective approaches to double-flow turboprop with afterburner weight estimation technique forming at the initial stage of its design, having potential of implementation with acceptable accuracy for new generation of engines.

The authors carried out analysis of the existing weight estimation techniques with different degree of their elementwise particularization, and under various methods of main regularities selection, linking engine gas-dynamic and weight parameters. “Modular” and integral engine weight estimation techniques were considered, and weights of 16 engines were computed using these techniques.

Based on carried out analysis, the sel ected variant of integral approach was updated with allowance for gathered statistical data on new generation of turboprop mass and gas-dynamic parameters. A correction factor, characterizing the generation to which a certain engine is related according to its weight efficiency, was determined.

Recommendations on weight estimation of an engine design based on the existing gas generator were developed. These recommendations imply implementation of correlation dependencies of the engine's separate modules weights fr om its operation parameters within the framework of the developed technique.

To determine the weight of turboprop with afterburner, developed on the basis of scaled or modified gas generator, a combined technique matching up either integral or “modular” approaches was formed.

Finally, the recommendations on implementation of the formed techniques with allowance for their future development by invoking additional data, including the data on newly developed engines, are provided.

In recent years, the problem of acoustic signature has become particularly actual and a topical due to the extensive use of combat aircraft systems with unmanned structures, solving decisive reconnaissance and strike tasks, for which low figures of acoustic signature ensuring is of prime importance.

The paper considers basic techniques for engine-propeller power plant noise reduction of aircraft type UAVs, including single air propellers of various structures and configuration, as well as piston engines.

Based on semi-empirical model the authors proposed equations allowing evaluate the effect of the diameter and number of blades on tonal components of the propeller noise in the condition of constant thrust, aerodynamic and geometric similarity of blade profiles, as well as the Mach number of the tip speed. Acoustic testing of Yak-18T light aircraft with two- and three-blade propellers, F30 and MAI-223M, performed at the Moscow Aviation Institute airfield, generally confirmed these equations qualitatively.

The propeller diameter decrease of a small-sized UAV with piston engine was considered as one of the options for noise and signature reduction. It was found, that the diameter decrease by 3.3% resulted in approximately 300 meters reduction of the distance to the ground checkpoint, which a small-sized UAV can approach without the possibility of being detected.

The features of acoustic pusher propellers and proposed methods for noise reduction are described. Based on the flight test the aircraft noise reduction afield technique by axial clearance increasing between the pusher propeller and the wing located in front of it was proposed. The paper demonstrates that with the considered clearance increase by an amount greater than the wing chord, the negative effect of the propeller mounting in pushing arrangement is practically eliminated.

UAVs designers can implement the engine-propeller power plant noise reduction methods, presented in the paper. Finally, the authors outlined the ways of further studies aimed at solving the problem of developing low-noise power plants for small-sized unmanned aerial vehicles.

The subject of the work consists in numerical study of the boundary layer on the wall of the combustion chamber and nozzle of a liquid rocket thruster. Using numerical integration method of the system of differential equations describing the boundary layer, the boundary layer parameters were computed as a function of the engine operating conditions and the pressure in the combustion chamber. To close the system of boundary layer equations, the values of turbulent moment and heat transfer coefficients are calculated by determining mixing length by the equation suggested by Prandtl with Van Driyst correction.

A numerical method for the boundary layer computation was realized as a software with the working interface in Excel. The program operates with relative dimensionless parameters.

Low-thrust LRE, burning such fuel components as nitrogen tetroxide and asymmetrical dimethyl hydrazine with the thrust of 200 N, parameters served as initial data for computation. The working flow parameters were taken according to the results of thermal and gas dynamics computation with average mixture ratio “on the wall” over the length of combustion chamber.

The paper presents computation results of the boundary layer parameters for the MAI-200-1 object engine: the displacement thickness, relative velocity profile, friction coefficient, nozzle flow rate.

The change of boundary layer thickness and flow rate coefficient for the object engine, and engines with working pressure of 2 and 3 MPa were calculated. The paper made clear that an increase in the combustion chamber pressure increases the relative thickness of the boundary layer, while nozzle flow rate falls.

The paper presents the results the experimental study of appropriately organized electric field effect, using electric unit for aviation kerosene impact (EUAKI), applied to kerosene flow at the nozzle inlet, on the kerosene-air mixture burnout temperature. TC-1 kerosene was used as hydrocarbon fuel. The air was fed to combustion chamber at the temperature of 150°C. Fire tests were carried out on the Samara State Aerospace University workbench.

Experiments on gas temperature at the outlet of combustion chamber gas collector characterization were performed by direct gas temperature measurement with single-point chromel-alumel thermocouple (operating temperature range from 0 to 1,100°C) shifted in the plane of the flow cross-section at the distance of 20 mm from the gas collector cutoff of combustion chamber combustor can.

The electric field parameters, such as voltage type at the EUAKI electrodes, its amplitude and frequency, and EUAKI design parameters effect on gas temperature distribution at the combustion chamber outlet while kerosene-air mixture burning. Atomizer modules herein, consisting of SPA “Salut” fuel atomizer itself and various EUAKI design with electric fields organization from different electric power supplies were varied.

It was demonstrated that implementation of EUAKI directly connected to the fuel atomizer inlet as a part of atomizer module by rubber hoses with corresponding permittivity increases the average and maximum gas temperature at the gas collector outlet up to 4.09% and 4.88% correspondingly, reduces gas temperature field non-uniformity at the combustion chamber outlet by 10.34% relative to the base.

The paper considers the solar thermal rocket engine (STRE) with isothermal (one-stage) and two-stage system concentrator-absorber system (CAS). It demonstrates their characteristics in the flight version, revealing rational parameters of the CASs under consideration, and inexpediency of attaining maximum possible hydrogen heating temperatures and maximal specific impulse with higher mirror booster accuracy in both structures.

For considered STRE schemes, implementation of heated hydrogen afterburning reveals the possibility of solar concentrator size reduction together with upper stage fuel compartment size reduction. Selection of expedient parameters of CASs under consideration may shift towards less accurate mirrors with less absorber heating temperatures followed by minor deterioration of upper stage ballistic characteristics.

To enhance STRE energy characteristics the authors suggest CAS with two-stage solar emission absorber, which heating level corresponds to the irradiance level in focal light spot. The highest hydrogen heating temperature occurs in the central part of the absorber. The specific impulse herein significantly exceeds the like when employing isothermal absorber.

Two-stage absorber efficiency computation regression model, based on energy balance of heating stages, allowing obtain rational temperatures relationship corresponding to maximum absorber efficiency, as well as optimal temperatures distribution along heating stages was developed. The obtained regression dependencies can be used for computation of real STRE, operating as a apart of space upper stage, flight characteristics. The paper demonstrates STRE flight characteristics with considered CASs, defines their specific flight It was demonstrated that in case of two-stage CAS mass efficiency exceeds the like for modern liquid means of interorbital transportation more than 2.3 times.

On oxidizer excess coefficient selection in case of hydrogen afterburning it is necessary take into account that for STRE with two-stage absorber each percent of concentrator diameter decrease corresponds to about one percent of payload weight reduction. This factor should be considered while practical design of various STRE structures.

The process of the liquid fuel atomization and vaporization is of fundamental importance for the GTE combustion chambers effective operation. Normally liquid fuels are insufficiently volatile, and therefore must be dispersed in large numbers of small droplets with an increased evaporation surface area required for the ignition process and combustion of the fuel-air mixture.

The paper presents the results of a new unique shadow particles anemometry method for studying parameters of the flame spraying nozzle unit of low emission combustion chamber (LECC) of the pneumatic type. A detailed description of the PSV measurement method and calculation algorithm when processing the data is presented. The special feature of this method consists in its relation to a method of direct measurement of various aerosols characteristics and provides highly accurate measurements of parameters compared to other methods. PSV method uniqueness consists in the fact that in addition to the spray basic parameters, it allows also define the shape of the particles, by freezing the shadows of droplets images in the measuring volume of camera matrix and high-speed pulsed backlighting. Tests were conducted on a CIAM laser diagnostics workbench in the open space behind the nozzle unit with fuel (kerosene TC-1) pneumo-spraying. During the tests distribution of fuel particles over size and shape at the distance of 30 mm from the nozzle section in the cross-section of spray pattern was obtained. Implementation of a new Shadow particles anemometry method (PSV) allowed verify experimental data, obtained earlier by the phase-Doppler anemometry, and the method itself has demonstrated its efficiency and effectiveness, as measured in terms of dense aerosols.

Topicality of stability and controllability characteristics selecting methods development, when employing rudder control, is caused by a number of incidents stipulated by the directional control channel drawbacks. An aircraft controllability in directional channel is completely defined by its dynamic characteristics, sensitivity and control efficiency characteristics, as well as by the characteristics defining yaw/roll motion interaction.

The paper considers earlier developed aircraft controllability criteria in directional control channel and analyzes them from the viewpoint of applicability to modern passenger plane with advanced automation.

One of the issues tackled in the paper consists in ensuring aircraft reasonable dynamic characteristics. All existing regulatory documents usually place demands on dynamic characteristics from the viewpoint of ensuring enough response speed in aircraft control channel. However, earlier studies revealed that unreasonably high response speed could become the reason of aircraft so-called «sharp response» on pilots effort. Thus, the requirements to should have upper bound. The paper presents the technique of criterion parameter determination, allowing determine an aircraft inclination to sharp response occurrence and the ways to its elimination by relevant selection of control system characteristics.

For modern aircraft with V-shaped wing and engines mounted on pylons, parameter, defining aircraft directional and lateral motions interaction, may attain rather high values. Automation introduction allows decrease this value, so that its equivalent value, i. e. the value with account for automation operation achieves an optimal value. The paper presents control system parameters selection technique ensuring optimal yaw/roll motion interaction.

The authors envisage two criteria to determine optimal control sensitivity. One criterion allows estimate sensitivity optimality in time domain, and the other in frequency domain. Both criteria give the same accuracy of the obtained results. The paper presents detailed technique for optimality evaluation of rudder control sensitivity in relation to aircraft dynamic characteristics and control stick loading characteristics.

The developed criteria give physical vindication of directional control channel characteristics optimality. They can be applied not only for preliminary selection of characteristics in directional control channel and ways of their realization on modern highly automated aircraft, but also for evaluation of mounted on the in-service aircraft.

An adequate external field of vision fr om the pilot's station is one of the topmost conditions of safe and comfortable aircraft control including a helicopter.

Explicit numerical values of vertical and horizontal vision angles from the main sighting point “C” are specified by regulatory documents, in particular by Aviation regulations (FAR-29). For clarity, the normative field of vision (FoV) is usually represented in the form of a chart  in rectangular axes, wh ere  и  are vertical and horizontal FoV angles respectively. The opening outlines should comply with normative chart as much as possible.

Currently used methods of view assessment, including a measuring method (natural and virtual with 3D model), are rather labor consuming, as they require human processing of measured data. Besides, with the initial data change, such as main sighting point “C” position, all the measurements must be repeated.

The author has developed an interactive technique of structural FoV plotting by means of Siemens NX8.5 – the basic 3D CAD system of the company. However, structural FoV does not take account for pilot's head mobility and human vision binocularity. It results in overestimated, sometimes impracticable, requirements for geometry of cockpit openings.

As a continuation of the above said research, the development of plotting technique for so-called actual FoV, complied with Standard 1 00444-81 and with due account of the above mentioned factors, has been carried out.

This problem was also solved by graphical method with Siemens NX8.5 CAD in a similar way as structural FoV chart plotting. As a result, actual FoV chart in rectangular coordinates has been obtained. All plotting, like structural FoV, are fully associative. With input data change, the geometry is reshaped automatically.

The author also managed to solve the problem of normative actual field-of-vision boundaries on a crew compartment surface, based on reverse combination of projections and convolutions of normative FoV boundaries in rectangular coordinate system. This allows optimize the location and form of cockpit openings at early design stages.

The paper deals with accuracy increase of uncontrollable aviation ground target killers implementation. It was found that the existing aiming systems use approximating dependencies in onboard ballistic algorithm, which does not allow provide high accuracy in all combat conditions due to introduction multitudes of assumptions.

As is well known, the kernel of a ballistic movement complete mathematical model consists in the system of twelve differential equations, which solution requires the set of means, ensuring its numerical functioning. These include equations describing ambient environment parameters, the system of inertial, traction and aerodynamic characteristics of ballistic objects, as well as data on initial and terminal conditions of thrown bodies' movement. Until recently, the low speed of computing facilities hampered with obtaining solution of uncontrolled air-launched weapons movement differential equations in the course of aiming. However, todays level of onboard digital technology allows overcome this shortcoming.

Therefore, in these conditions we have the possibility to realize the onboard ballistic algorithm based on numerical solution of differential equations directly onboard an aircraft in the course of aiming.

The authors analyzed the ballistic problems solution accuracy during modern aiming systems terrestrial fire, implementing approximating dependencies in onboard ballistic algorithm, and revealed their main shortcoming, namely, impossibility of ensuring high accuracy of application in all conditions of combat operation, as well as with various operating lives. New technique and ballistic support algorithm for aviation uncontrolled destruction facilities were developed based on complete ballistic movement model solution. It allowed enhance the range of tactical employment due to firing initial conditions definition, atmospheric parameters, and aviation destruction facilities movement trajectories parameters definition; nutation angle prediction; flight trajectory parameters introduction into aiming system, and angular correction computation for aviation artillery-type weapon and uncontrolled aviation missiles, with allowance for the predicted nutation angle.

Notwithstanding the great number of publications concerning pilot models, none of them considers the issue of stick force optimal characteristics selection. It can be explained by the fact that pilot describing function in visual signal tracking is insensitive to loading characteristics variation and does not allow reveal any regularities and their effect on pilot model parameters.

This paper is aimed at studying the effects of stick force characteristics on pilot model and its components, such as limb-manipulator and neuromuscular systems, as well as finding objective proof of loading characteristics, selected by pilots, optimality.

The paper presents recently obtained experimental data on the effect of control stick force characteristics, such as gradient of stick-force damping on pilot model parameters. The effect is analyzed based on pilot model frequency response identified in the problem of compensatory pitch motion tracking. For limb-stick and neuromuscular systems characteristics identification, input strain signal is introduced in addition to visual input signal. Frequency response characteristic computation of various pilot model components was made according to specially developed program, based on fast Fourier transform.

Analysis revealed that the force gradient variation affects neuromuscular frequency response, demonstrating thereby a pilot's adaptation to the stick force variations. Due to this, the limb-stick cutoff frequency of the open-loop system remains constant for the force gradients assessed by the pilot as optimal. The force damping does not have any significant effect on limb-stick system frequency response.

The obtained results are of regular character and contribute to theoretical and practical aspects of pilot models implementation for aircraft sensitivity evaluation.

The paper describes the process of spatial movement of the aircraft relative to the earth coordinate system by a system of differential equations, taking into account the dynamics and kinematics of translational and rotational motion.

The aerodynamic impact of the environment on the aircraft is determined by its configuration, position of the associated coordinate system relative to the velocity of the aircraft center of mass and vector of its angular velocity. To ensure an aircraft steady state flight mode the model solved the problem of balancing, consisting in the engine thrust values, angle of attack and the deviation of the aircraft control organs selection with subsequent solution of the system of differential equations. The output variables of the model are the parameters characterizing the actual position of the aircraft in space.

The calculated missile trajectory, represented in the form of differential equations and algebraic dependencies, describes the missile guidance to the aircraft. The result is the relative distance value of the aircraft defined by the elevation angles and azimuth. The rocket direction of motion measurement is made according to the method of proportional guidance. The control system sets the missile maneuver with an overload, directly proportional to the angular velocity of the rocket-target line of sight. Thus, any time it tends to ensure the direction of the missile movement to the set-forward point.

Based on the canonical equation of motion of the center of mass of the active protection element the terms of its ejection, to deliver it to the point of space where the guided missile is situated was calculated. By simulating the flow over the active element by turbulent incident flow, using finite volume method in Ansys CFX the authors defined the ballistic coefficient of the active protection element. It allowed us to calculate the resistance function value and produce the data on trajectory and projection parameters of active protection element to the control unit.

As a result, this model allows calculating, under different tactical actions of the aircraft crew, the target miss, the orientation angles and missile speed of convergence with the aircraft. When processing simulation results one can obtain the characteristics of the missile encounter with an aircraft, as well as active element ejection parameters for its encounter with the rocket in calculated space point.

The paper deals with project quality management in aerospace industry. It analyses acting domestic and foreign Standards on project management, and offers problems grouping for promotion in the field of innovative technological projects aimed at aerospace products development quality management.

The paper suggest to eliminate technological gap between Russia and European and American industrial enterprises, developing samples of advanced technology in the sphere of aviation and astronautics through implementation of practices accumulated in international Standards on separate projects management, as well as programs and portfolios managing. Innovation projects management quality increasing for National design departments should base on quality provision and management integrated system development and putting it into practice based on both ISO Standards and proper Projects and Programs management Standards. The existing classification of design performance and other indices needs to be improved.

The author envisages concepts of technology, technological innovation project and technologies transfer. Depending on complexity, technology can be included into economic turn-round. Transfer key criterion means technology working efficiency in terms of technological complexity. The State puts forward the problem of research carrying out on improving the system of innovative projects in the area of technological processes provision and management. Its solution options are significant of applied research planning procedure renewal, and Hi-Tech innovation projects realization technological provision program substantiation at the life cycle early stages, and innovative constructive-technological solutions marketability preliminary estimation, as well as process design planning optimization based on Hi-Tech projects with allowance for economical production.

The results of the study can be implemented for new Hi-Tech innovation projects management quality methodology development.

The paper practical has practical importance for acting quality management system at the aerospace industry enterprises improvement. It can be implemented also in the process of specialists training in the innovation projects management sphere.

The paper envisages circuit solutions for transformer multifunction switched mode converters meant for uninterruptible power sources as constituent parts of onboard aerospace electric power facilities and electric power supply systems. All solutions are protected by the Russian Federation priority. The paper is of interest for a wide range of specialists working in the field of aerospace onboard power electronic equipment design.

To power actuating brushless motors (aircraft onboard equipment in particular) the voltage higher than that provided by batteries, solar of supercapacitor (ionistor) elements is required due to the necessity of varying magnetic field space forming (either circular or linear) by currents flowing through flexible wires. Ensuring relatively higher voltage level only through series (stacked) low-voltage units, or series connection of the above said batteries with significant supply currents is hampered technologically, and leads to mass and size, reliability and cost parameters reduction. For example, in case of “stacked” units leads burning-out (or break) they should be shunted by diodes with low-voltage conducting junctions. In case of breakdown at the leads of a parallel link, it should be provided with disconnecting fuses. Thus, to increase the voltage level of a primary relatively low-voltage source switched mode converters (SMC or DC/DC converters) based on field-effect transistor switches (MOSFET) with low Rds(on) should be used. They should herewith be reversible to provide feeding batteries intensive charging. As a rule, such converters are included in so-called secondary power sources, or stand-by uninterruptible power sources (UPS) fed by batteries [1].

Aerospace uninterruptible power sources included in onboard electric power facilities and electric power supply systems, acquire primary energy from chemical or solar batteries, either form newly developed super capacitor (ionistor) batteries with relatively low voltage (28 V). As a rule, the UPS output voltages herewith are higher DC voltages (such as 135 V, 270 V, or 540 V), or higher AC three-phase (or single phase) voltage (stabilized or regulated) of constant or regulated frequency (e. g., within the limits of 115/200 V, 360-800 Hz, or 0-115 V, 0-400 Hz). Besides, UPS should provide fast feeding battery charging (accumulator or supercapacitor).

In this regard, at least specific requirements are placed on the above mentioned UPSs.

This article describes one of the powder metallurgy methods, namely mechanical alloying (MA), used for composite materials production. MA is a solid-phase process of deformation impact on the powder material. MA changes the structure and properties of the processed materials. The authors analyzed the effect of technological modes on the process of mechanical alloying. They described, in particular, the main types of ball mills, employed for МА process carrying out. The authors examined the effect of the impurities on fractional, chemical and phase composition of composite granules, which can both accelerate supersaturated solid solutions and amorphous phases building-up process, and prevent diffusion to form amorphous oxides and phases with work material. The authors demonstrated in the paper that the shape of the shape of the container and grinding bodies could also affect the MA process and its results, as well as MA effectiveness and fractional composition in particular. Shape, size and material of the grinding bodies selection depends on several factors. Generally speaking, the grinding bodies should correspond to two basic requirements, namely, they should possess developed superficial area to provide contact with the processed material, and have enough weight to possess enough energy for processed particles grinding. The grinding media can be not only in the form of a globe, but also cylindrical et. On the Al-50% Ta system example the authors envisaged the effect of globes weight to the weight of a material ratio on the MA process.

The authors demonstrated also that the MA rate is one of the most important parameters affecting the process of the processed material grains mixing and grinding, chemical reactions process and phase transformations occurring in solid phase. It is well known, that the greater the mill rotation speed, the greater the kinetic energy transferred to the bodies and particles, and, hence, the intensity of the process increased. However, excessively high rates might cause a number of complications, such as grinding bodies high degree abrading and overheating either of a drum mill, of processed material. The authors also studied the issue of temperature effect on phase and structural transformations during technological process. They noted, that high temperature contributes to phase transitions and chemical interaction, while lower temperature works towards nanocrystalline state and metastable phases forming, as well as allows process plastic materials effectively.

Thus, the materials presented in the paper help not only to select the initial charge materials processing mode, but also predict the obtained results.

Titanium intermetallic Ti₂AlNb (orthorhombic phase) based allows are promising materials for gas-turbine engine elements manufacturing operating at the temperatures of 650 — 700°С instead of heat-resistant steel due to their high specific properties, and also intermetallic super- - and  -alloys possessing low technological plasticity.

Orthorhombic alloys phase composition and structure strongly affect the final mechanical and technological properties of semi-finished products, which can be controlled by certain of thermo-mechanical and thermal treatment modes. Thus, the purpose of this study consisted in studying the effect of heat treatment on the structure and properties of heat-resistant alloy based on Ti₂AlNb titanium intermetallic.

In this work, the effect of various heat-treatment modes on the structure, hardness and mechanical properties of the VTI-4 alloy based on Ti₂AlNb titanium aluminide was studied. The samples were subjected to heat treatment, X-ray diffraction and metallographic analyses. Besides, the hardness of samples was measured by Rockwell method, and mechanical tensile tests were carried out at room temperature.

Based on conducted studies, data on the temperature ranges of phase regions in the alloy was obtained, and a scheme for the two-stage heat treatment was designed. It was stated, that the structure and hardness of the alloy are greatly affected by the cooling rate between the first (high-temperature) and the second (low-temperature) treatment stages. Increasing of the cooling rate from 0.01 K/sec to 10 K/sec resulted in fine-dispersed orthorhombic phase formation; the alloy hardness increased by 5 HRC units, and the strength grew by 100 MPa while maintaining a satisfactory level of 4 — 6% for the plastic properties. The cooling rate after the low-temperature stage had no effect on the alloy structure and hardness.

It was shown also, that temperature reduction of isothermal holding in the low-temperature stage by 50°C resulted in the tensile strength increase by 80 MPa, and plasticity decrease by 3%.

Designed VTI-4 alloy heat treatment modes on the example of rod semi-finished product allowed form in the alloy structure with different size of structural components. The obtained results allow also predict changes in the strength and plastic properties of other types of VTI-4 alloy semi-finished products according to the need for further forming operations.

The gas turbine engine advancement goes hand in hand with the development of its basic component, namely, gas turbine as the key source of efficiency enhancement of the engine in aggregate. With each turn of gas turbine development, materials and technologies used for its manufacturing became more and more complicated and, as a consequence, expensive. Russia is one of global manufacturers of gas turbine engines.

The cost of engines for aviation and power industry applications is considerably high. Thus, on this background its reduction remains the main criterion of manufacturer's competiveness on the market. Besides, we should bear in mind that the gas turbine engines maintenance costs in the course of the engine life might exceed its original cost. Without effective maintenance technologies, manufacturing would incur permanent losses. One of the basic specifics of gas turbine engines consists in their significantly high production costs of a number of their parts and subassemblies with relatively short lifetime, requiring permanent replacement. Rotor blades present precisely these parts. They can be damaged by a great number of factors from changes in the structure to loss of geometry. The latter is the most frequent factor even in the case of insignificant geometry loss. From the maintenance technologies viewpoint turbine blades restoration is the most cost-effective, compared to the other parts of the engine. But the complexity of this task remains the major obstacle to its realization.

This article discusses the possibility of using high-temperature solder powders as wear-resistant layers applied by laser pulse buildup, as an alternative to classic wear-resistant composites with tungsten carbide admixture. These materials are undergoing testing for further pen height recovery on the example of the turbine blade of the turboprop starter for NK-12MP aircraft engine, and attaching wear-resistant to its end edge. Based on the conducted studies with Tescan VEGA3 LM electron microscope and Hardness DuraScan-10 micro-hardness meter, together with local abrasive wear tests and various powder materials, such as VPr11-40N, VPr24, VPr27 Rock-Dur 6740, analysis while pulse laser powder buildup, the authors confirmed the applicability of several solder powders as wear-resistant layers for turbine blades contact surfaces recovery. Further, comparative studies of the basic material, soldered and built-up structures of VPr11-40N (having the best figures) solder were conducted to detect hardening wear-resistant phases. The cooling rate dependencies of shaping and VPr11-40N solder strengthening phase size were revealed.

Ensuring competitiveness of the Russian Federation in the conditions of strengthening of the global competition defines the innovation development of national economy as the priority direction. This direction realization assumes dynamic and intensive development of the industry basis on development and implementation of radical, cardinal, breakthrough innovations, and primarily technology and product. This, in in its turn, imposes increasing requirements to research and manufacturing base of the Russian industry.

Development and upgrading efficiency, productivity of research and manufacturing base of industry represents difficult, complex, coordinated process of its participants' interaction. They involve all the basic, vital and concerned parties, and are aimed at innovative cycle reduction, primarily, at the stages of innovation development and innovation activity growth, and finally, holding leading global positions over key, priority trends of technological development.

Realization of this process in the framework of the Russian Federation of a scientific and technology complex development assumes continuous improvement of its organizational and economic mechanism. One of the main methods providing development of a scientific and technology structure of Russia consists in scientific and technical projects realization within the framework of the State order, requiring forming and carrying out competitive selections of such projects.

The solution of the above-stated problem proposed in this paper consists in forming specialized organizational and executive structure of a project realization within the system integration of participants-contractors working on the project.

This model of forming organizational and executive structure of the project is developed based on the criteria accounting for extra income on the project, possible from implementation of collateral and intermediate product output, while developing research and technology reserve and, thus, under otherwise equal conditions, supplements the project economy and reduces the risks in case of possible losses.

The airlines employ flexible approach to strategic planning and conducting regular environmental monitoring of an unstable economy. The purpose of this paper was to adapt the SWOT-analysis to assess regional aviation clusters development.

This cluster incorporates enterprises, interconnected by the aircraft lifecycle: aviation enterprises, infrastructure enterprises, operating airlines, scientific and educational organizations. The method of the SWOT-analysis is applied for a specific enterprise and for the aviation cluster as a whole. The author analyzed the factors of the external and internal environment of the Ulyanovsk region aviation cluster.

In the studying process the priority of strategic decisions and the most significant capabilities of this socio-economic system, presented the basic economy indicators of this cluster was identified.

The study proved that using the SWOT-analysis requires quantification and ranking of external and internal environment factors of the regional aviation cluster.

Quantitative correlation of the factors reveals negative phenomena in the external environment of the aviation cluster. The paper offers the expert evaluation of factors according to four criteria such as, factors rating calculations, the relationship of the identified external and internal environment factors of the aviation cluster.

The greatest threats for the Ulyanovsk region aviation cluster are as follows: the gross regional product decline, the risk of the growth rates of loans to airlines, severer tax environment for business, dependence of the airlines development from State support. Risks that occur quickly and unexpectedly, devastating to the economic system. If adverse factors are inevitable, but not instantaneous (for example, the outflow of the region qualified personnel, reduction of the population), the production is adapting to them.

The research has practical value due to quantitative justification of the priority risks enabling the company to direct the limited resources more precisely.

The important role in development stability and efficiency enhancement of national economy belongs to small and medium business (SMB), as an echelon of economic dynamics. This is confirmed by the experience of economically developed countries, where the share of small businesses constitutes 56% of GDP. It should be noted that the of SMEs successful development in these countries became possible due to active government support (tax, legal, organizational, personnel, etc.). Unfortunately, in Russia small business is developing slowly, and one of the important reasons consists in the the lack of systematic State support. Today, in the conditions of economic crisis caused by the endless sanctions, special attention should be given to small businesses in the manufacturing industries, especially in machine-building industry. All the more so, in these industries, small business accounts for only 15-16% of all active small businesses. It is important to note, that development in these sectors should be based on close cooperation with large corporations. In this situation, small business can take the risk of the releasing new prototypes of high-tech industrial products. They can also take over the production of components for large enterprises, thereby reducing costs. Organization of small businesses in these sectors requires a fundamentally new approach. At the stage of economy modernization the SMBs need a more sophisticated system of Government support, which should include: development of programs for the development of SMBs cooperative relationships with large manufacturing structures, creating conditions for access to the scientific and technological achievements; assistance in professional staff training and retraining. Thus, for radical strengthening of small and medium business role in manufacturing industry it is necessary to develop a fundamentally new strategy of state support, that will contribute to the development of the organization and management of innovative-investment activities in small business.

Cluster policy has become the dominant trend in the development of many regions. The modern model of aviation industry enterprises consolidation assumes parceling of enterprises on several levels, such as suppliers of raw materials; suppliers of parts and components; suppliers of components and assemblies; sub-integrator; final integrator. Aerospace Cluster of Samara region is one of the high-tech sectors of regional economy. Bearing manufacturing industry seems to be interesting and prospective, since they present the represent components of every rotating mechanism, implemented in every branch of production (including aircraft and helicopter). In addition, products of the plants of the branch under consideration is required either in the region, in the country or other countries.

Currently, due to the marked production growth, produced products nomenclature increase and expansion of sales network enhanced the role of logistics significantly. Competent organization of logistics operations allows obtain quite considerable cost savings, which is an important tool for industrial enterprises production and commercial management activities.

Sales revenue from products sales factor analysis over three indicators, namely, product unit annual average cost; annual production output, which, in its turn, depends on the staff on the payroll and annual average yield by a single worker. The results of the analysis allowed make a conclusion on the necessity of transportation process optimization, since transportation costs constitute significant share of product cost and final product price.

In view of wide geography of sales, the decision was made on the necessity for establishing a distribution warehouse. In the course of calculations, the optimal warehouse location based on cities remoteness from a production point and their annual claims, was determined, and transport selection for production delivery was made. The structure of intracity production distribution on the example of Samara was offered, and the travelling salesman problem was solved, using the two-parameters accrual method, namely time and distance.

The study bears the applied nature, and the work has practical value when minimizing transportation costs and embodiment of transport and logistic activities, which will lead to effectiveness enhancement of the industrial enterprise.

The study is an applied nature, and the work is of practical value while minimizing the costs of the transport and logistics activities, which will totally increase the efficiency of the entire industrial enterprise as a whole.

A serious problem of modernity is a problem of flight safety promotion. This problem needs to be addressed not only during operation but also during the design and manufacturing of aircraft.

With this objection in mind, it is necessary to ensure implementation of Aviation Activities Safety Management System at the enterprises of aviation industry complying with the State Standard (GOST R 55848-2013), System of Safety management (GOST R 55585-2013), Quality Management System (GOST R ISO 9001:2015), Environmental Management System (GOST R ISO 14001:2007) and Occupational Safety and Health Care Management System (GOST R 54934-2012/OHSAS 18001-2007).

Currently, many aviation enterprises are putting into practice the system of Quality Management, but Environmental Management System and Occupational Safety and Health Care Management System are not so actively introduced in industrial enterprise management practice, notwithstanding that Russias annual underproduction due to industrial diseases and injuries goes as far as one trillion rubles.

Implementation of Occupational Safety and Health Care Management System will allow reduce these huge losses.

With implementation of Occupational Safety and Health Care Management System aviation enterprises acquire real economic effects by improving working conditions, reducing the lost work time as a result of injury and disease, reducing costs of benefits andcompensation for work in harmful working conditions, improve labor productivity and production growth.

The equations for economic effect and efficiency of Occupational Safety and Health Care Management System calculation demonstrate that the enterprise acquires not only social, but also real economic effect and social and economic efficiency.

The proposed equations are recommended for implementations not only for computing the economic impact and effectiveness derived from the introduction of Occupational Safety and Health Care Management System, but also for management decisions related to the implementation at the enterprises of aviation industry.

Possibility of control by aerodynamic characteristics of a large aspect ratio elastic wing with the end winglets turning in the wing plane is investigated.

Controlled twisting of the elastic wing in flight subjected to aerodynamic load which depends on the wing twisting can be carried out by turning of small end winglets in the wing plane with the help of a small power drive.

The coupled aeroelasticity problem is solved using mathematical model based on the discrete vortex method for calculation of aerodynamic loads on deformable wing and the wing as a thin-walled weackly-conical beam subjected to bending, transverse shear and torsion.

The numerical solution of the aeroelasticity problem is obtained for the large aspect ratio wing with the winglets turning symmetrically forward or backwards in the wing plane. Due to turning of the winglets there appear the end aerodynamic moments which cause elastic twisting of the wing and change the distribution of the aerodynamic load along the wing.

For the example of a rectangular wing with the winglets it is shown that the turning of the winglets it is shown that the turning of the winglets in the wing plane creates the end torques and significant angles of twisting of the large aspect ratio wing and as a consequence significant change of the aerodynamic loads and the wing aerodynamic characteristics.

The results of calculation show that in a case of a wing which is sufficiently pliant in twisting in the wing plane at the angle δ can be effective for control of the wing aerodynamic characteristics . In case of a wing which is sufficiently rigid in twisting the winglets become ineffective.

The presented work is dedicated to horizontal empennage multidisciplinary optimization method development. Horizontal empennage is a complex technical system, described by the equations belonged to various scientific disciplines. That is why the developed method is called multidisciplinary. The horizontal empennage efficiency can be evaluated by the values of generated pitch moment and its gradient, guaranteeing the aircraft balancing and specified flight maneuver execution. The object region analysis was carried out and various parameters combinations for optimization within the framework of the given problem were determined. We determine optimization line and specify weighting factor for each parameter. Each of the parameters can be set either as a range-parameter, fixed-value, or a set of discrete values. Besides, the simultaneous several parameter setting by a set of tuples, containing discrete parameters values, is possible. The goal function is obtained (where the number of addends is determined by the number of optimized parameters). The goal function structure allows operate with all optimized parameters combinations, regardless of the way of their specifying.

Various approaches to the horizontal empennage optimization (methods employing the Pareto principle, and the Hurwitz criterion) were studied. The analysis of the obtained results revealed the insufficient efficiency of the implemented methods. To improve the obtained results, a new multidisciplinary optimization method was developed and suggested. This method employs several evaluation functions to obtain optimal solution. The efficiency of this method is demonstrated using various data sets and input data combinations. The effect of various weighting factors values on the obtained result was studied. The result of the suggested method implementation is horizontal empennage geometrics.

The suggested method was realized in the form of a Runtime library and integrated into CAD system Siemens NX 7.5 “Modeling” environment.

The paper presents the results of computational, experimental and analytical studies of gas-dynamic processes in Ranque-Hilsch vortex tubes. The presented review considers the relevance and need for employing vortex effect for aerospace engineering. It reveals the necessity for vortex tubes with varying geometrical dimensions design for the purpose of operation range enhancement. The authors developed vortex pipe 3D model in SolidWorks system. They realized a viscous gas in vortex tube computation engineering method, and demonstrated its implementation results in gas-dynamics computing FlowSimulation pack. To solve this problem Reynolds averaged Navier-Stokes system (RANS) of equations was used in this work. All computations were performed with orthogonal computing net using finite volume method. Two-parameter model of κ — ε type allowing sufficient flow core resolution was used as turbulence model. Several basic vortex effects, such as injection, heat stratification and vortex inversion, were obtained by computation. All calculation were performed for various structural versions. A series of experiments was conducted with custom-made experimental setup. Processing of the obtained results lead to obtaining hot and cold flows productivity optimums, injection ratio, temperature stratification, as well as adiabatic and temperature efficiency.

The experimental results fully confirmed the vortex effects of obtained by engineering computational method. The authors suggest new differential equations for parameters computation in these tubes. The obtained equations establish relation with flow rotation and whirling, as well as explain the enthalpy effect. Computational and experimental as well as analytical studies should continue with regard to optimal structural concept.

The extensive development of small and unmanned aircraft together with existed requirements to permissible levels of noise generated by aircraft, make the noise prediction problem afield topical for prospective aircraft with engine-propeller power plant. The main source of noise afield created by aircraft of such kind is a power plant, consisting of single propellers of various design and configuration, and piston engines.

This work integrates and develops the authors’ previously developed methods of computing separate propeller noise and the piston engine noise for solving the problem of forecasting the characteristics of light aircraft and unmanned aerial vehicles power plants’ total acoustic field.

The authors suggest a semi empirical model for noise levels, generated by aircraft piston engines in the far field, evaluation with allowance for main noise sources. The acoustic field is considered as a superposition of fields, formed by propeller and piston engine noise radiation. For propeller audio frequency levels estimated evaluation implementation of semi empirical method developed earlier by the authors is recommended. To determine propeller’s vortex noise levels, presumably dominant in the broadband noise of tractor propellers, we propose to use one of analytical models of the trailing edge noise. To calculate the acoustic performance of the piston engine we suggest to use an empirical noise model.

The paper demonstrates close agreement between computed and experimental data on power plants with tractor propellers. Experimental data on power plants noise was obtained during light aircraft of An-2, Yak-18T, MAI-223M and F30 acoustic trials under static conditions at the Moscow Aviation Institute airbase. The acoustic field herewith was supposed axisymmetric relative to the propeller axis, while test microphones were located at the ground level. It allowed exclude the interference of sound impact on measured noise levels.

The future trends of the study concerning improvement of the above mentioned method and extension the area of its application were formulated.

The paper considers the issue of low-thrust liquid engine powered by nitrogen tetroxide and dimethyl hydrazine components non-symmetrical dimethyl hydrazine thermal state theoretical study with account for boundary cooling. The goal of the paper consists in analyzing the results of combustion chamber wall thermal state computation at various operating modes, such as steady-state continuous mode with stationary and non-stationary thermal field, as well as steady-state pulse mode.

Liquid rocket engine MAI-200-1 developed in the laboratory of MAI “Liquid rocket thrusters” and undergone fire tes sel ected as a subject of research.

For thermal state computation, the authors used mathematical model based on the proposition of combustion chamber wall incoming and outgoing heat flows equality. To solve non-stationary heat problem the differential Fourier-Kirchhoff heat equation in cylindrical coordinates in the case of stationary environment and the absence of internal heat sources is used. Pulse mode of the engine operation is modeled by a quasi-steady approach when non-stationary modes during engine starting and voiding are replaced by the set of stationary modes with intermediate parameters.

Oxidizer and fuel were considered as boundary cooling components to protect the combustion chamber walls fr om hot combustion products impact.

Computation results prove selection of fuel as boundary cooling component with relative boundary mass-flow rate not less than 20%. Under these engine operating conditions it will allow sustaining the wall temperature within the limits of maximum permissible temperature for ХН60ВТ material.

The combustion chamber wall thermal state for pulse operating mode with various on-time and off-time values, such as on-time of 1 s, off-time of 1 s and on-time of 0.05 s, off-time of 0.05 s were analyzed.

Presented computation results may be interesting for specialists working in the field of liquid-propellant thrusters, as well as for specialists occupied with spacecraft propulsion systems design.

The purpose of this work consisted in determining the type and functional content of the dependencies, constituting the two-component LPRE control algorithm and obtaining formal description of these dependencies for further use of this algorithm while implementing the engine as a part of a launcher during the flight.

It is shown that the task of maintaining the specified for flight conditions engine thrust level values R and mixture components ratio Km are clearly described by specifying functions of regulator assembly drives position in relation to the six parameters: R, Km and four conditions at the engine inlet (temperature and fuel components pressure).

This conclusion was drawn by analyzing the structure and functional dependencies of LPRE mathematical model. It was successfully proved by determining such dependencies using adequate fire tests results of a given single-chamber LPRE approximation.

To determine control algorithm for LPRE, undergone hot testing, the author suggested implementation of computational-experimental model (CEM), formed according to the results of this engine hot testing.

The properties of such model allow carrying out reliable forecast computations of the engine operating procedures parameters in a wide range of the six parameters under consideration, namely operating modes and ambient conditions.

The final form of control algorithm represents a polynomial, approximating computation results based on CEM, carried out over six-dimensional array of computed points, defined within the required engine operation range.

The adequacy of the proposed approach to the control algorithm formulation in the wide range of all six parameters is validated by comparing the values obtained by approximation with experimental data of a given single-chamber LPRE.

Heat and hydrodynamic processes are becoming determinant while creating new types of engines for space, aviation and nuclear power systems [1 – 9]. Unsteady hydrodynamic and heat transfer processes study is an extremely important problem of engine building.

Only the combination of fundamental and engineering studies provides most effective way to design precise unsteady process model for practical computation. Experimental studies carried out in Moscow Aviation Institute (MAI) hold a prominent place in this field [10 – 17].

The turbulent flow structure studies carried out in MAI reveal non-stationary conditions fundamental effect on turbulent flow structure.

Axial and radial velocity and temperature pulsations, average parameters and their correlations were measured as a part of the study. Generalized experimental data reveals significant impact of flow acceleration and deceleration on turbulent structure. Three specific zones in turbulent flow were identified: near wall area y/R = = 0...0.02 (y — distance from the wall, R — radius of the channel); maximal turbulent parameters modification area y/R = 0.02...0.4 and flow core. Significant difference of turbulent viscosity between steady and unsteady approaches up to three times was identified. Comparison of quasi-steady and unsteady approach to heat transfer and hydraulic resistance coefficients revealed the two-times difference. Undoubtedly, such huge difference is unacceptable for space, aviation and nuclear energetics. This result agrees well with experimental data obtained by other authors [18, 19].

Based on non-stationary conditions significant impact on turbulent structure a computation model was developed. With flow acceleration, hydraulic resistance coefficient exceeds relative quasi-steady value by 2 times and more. During flow deceleration, it is 35% less.

Experimental study results present reliable base for further theoretical studies to be carried out in MAI [17]. The existing high-Reynolds turbulent models are not able, in principal, to consider non-stationary effect. From turbulence models analyzed in [18], only Menters SST model, which is low-Reynolds model, gives the results close to the experimental. Generalized equations for non-stationary friction and heat transfer coefficients at flow acceleration and deceleration in a tube for engineering design were obtained. The advantage of such models consists in the possibility of their employing for any monotonous flow variation curve, as well as satisfactory convergence with experimental data on hydrodynamic non-stationary gas flow in through channels [20].

Among the works of theoretical character, the studies of Professor Igor Derevich should be noted in the first place. In reference [21] the author considers the gas flow with monotonous consumption decrease/increase, and reveals the causes of computation and experimental data mismatch.

For practice, we recommend to analyze the effect of non-stationary processes on a certain jet engine control system. In case, when the processes are principally non-stationary and the required accuracy must be high, a non-stationary model and/or other approaches, considering non-stationaries, should be used.

High-speed transport design, aircraft engines ecology and higher energy efficiency guarantee by improving fuel atomization and air-kerosene mixture quality in aircraft engines intensive research is carried out. To improve fuel atomization and air-kerosene mixture burning we suggest the use properly shaped electric fields in atomizer fuel supplying contours. For the first time the authors studied the effect of variable frequency AC electric field on combustion products chemical composition, when employing kerosene TS-1. Experimental results on the effect of variable frequency AC electric field on air-kerosene mixture combustion products burning rate were presented for the first time.

Post-combustion flow speed measuring at the simulative combustion chamber outlet were carried out at Samara State Aero-space University (SSAU).

Air-fuel mixture combustion products burning speed experimental determination technique was developed at SSAU. It forms the basis of the research on the effect of AC electric field on air-kerosene combustion mixture products speeds.

Employing the speed measuring data, computations of superficial velocity and mass flow ratio were carried out using well-known equations for gas-dynamic functions.

The result of experimentation consists in creating Microsoft Access database file with further possible export to Excel.

Experimental studies were carried out at SSAU on a single-burner bay of a simulative combustion chamber with operational OJSC “Klimov” duplex nozzle for liquid fuel. We employed a swirler with blades angle φ = 72°10′; gas collector with cone outlet diameter of 133 mm; square spacer plate with square cross-section shaped with square side of 180 mm and a baseline case of offset area holes, when mixer apertures were open. Kerosene TS-1 was used as fuel. Low-pressure compressed air was fed under pressure ≤ 0.75 MPa, and solid tracing particles were used for laser measurements of Ch-4 type.

When the AC electric field was applied to kerosene along each diameter, prior to feeding to atomizer, speed values move intermittently up and down. With this, air-kerosene mixture combustion products maximum relative speed reduction was 2.45%, while maximum relative speed of air-kerosene mixture combustion products with applied to kerosene flow AC electric field at the outlet of combustion chamber was 1.425.

Propellers design and development for modern coaxial propfans and their automated control systems are impossible without in-line simulation test benches, which allow reduce testing fee, imitate failure situations, work through control laws and algorithms and determine automated control systems stability margins.

Turboprop engine mathematical model plays key role while testing propellers and automatic control systems with in-line simulation test benches. The tests validity depends on accuracy of non-stationary processes reproduction by mathematical model. Due to turboprop dynamic characteristics errors when employing linear methods of modeling, at present, non-linear element-by-element models became widely used. In the course of SV-27 coaxial propfan and RSV-27 hydro- mechanical regulator testing bench, JSC SPE “Aerosila” employs D-27 turboprop non-linear element-by-element model. Implementation of gas turbine engines non-linear models results in significant processing power waste due to the multiple recalculation of the thermodynamic mathematical model while compressors and turbines joint operation point search. To optimize the computational process while using a non-linear turboprop engine mathematical model the authors suggest to use of a genetic algorithm. Genetic algorithm was developed with LabView software, employed with in-line simulation test bench and associated with the engine mathematical model. Genetic algorithm of various configurations and probability values of mutations and number of species in population with in-line simulation test implementation efficiency was studied. The results of the study allowed determine the optimal genetic algorithm configuration and parameters of its optimal operation. In its optimal configuration with a small number of species in population and increased calculating error, this genetic algorithm appeared to be effectiveness comparable to method of successive approximations by bisection. However, the genetic algorithm execution instability, leading to computational resources wasting for some calculated points, makes its implementation in turboprop engine mathematical model, used with in-line simulation test bench for air propellers tests and their automated control systems, impractical.

The main objective of the research, aimed at developing combustors for civil aviation, consists in ensuring competitive level of engines emission characteristics. The presented work is dedicated to the development of technology for uniform air-fuel mixture preparation in the flame tube head with respect to aircraft combustion chamber.

Gas turbine engine aggregate characteristic guarantee, such as reliable start-up, wide range of stable operation, fuel combustion efficiency and low noxious emission depend in particular on combustor reliable operation. The researches in this field for the most part are agreed herein that achieving high-level of the above said characteristics in the combustor is stipulated, not after all the others, by liquid fuel crushing process quality and its preliminary mixing with air in the flame tube head. It is known that combustion of previously prepared homogeneous air-fuel mixture in model heat generators allows obtain low outlet noxious emission. However, real GTE combustor has no place or the time for such preparation. It stands to reason, that it is necessary the employ the available space and the residence time maximally to direct the air-fuel mixture characteristics drift towards a homogeneous composition.

This work presents the results of the designed flame tube head with liquid fuel pneumatic atomizer for low-emission combustor. The paper describes the air-fuel uniform mixture preparation technique in the flame tube head with fine-dispersed spray in swirl flow conditions.

Autonomous tests of the developed the flame tube head have been conducted. In the course of these tests the main characteristics of the air-fuel spray formed after burner by a non-contact laser diagnostics method in open space conditions were studied. According to the results of cold tests, the average Zauter diameter of the fuel droplets in the idle mode is about 23 microns. The wide and intense backflow zone is formed near the device axis. To test the developed device and method of air -fuel mixture preparation, fire tests in the model three-burner compartment under high-pressure environment were carried out. The ignition and blowout points under earth conditions have been obtained as the result of tests conducted. The efficiency of lean air-fuel mixture combustion technology has been confirmed.

Revealing climatic conditions effect in the course of pre-launch procedure of reusable space system is a necessary condition of thermal protection structural components ground development. The climatic tests experiment must simulate positive and negative temperature effects, as well as humidity and precipitation effects. The climatic tests algorithm is designed as a unified processing chain of test set up when a number of simultaneous or consecutive experiments are conducted at the experimental testing bench. The algorithm includes thermal-stability, low-temperature stability, moisture and weather resistance tests. The presented paper describes methods of reusable thermal-protective materials structure testing under the above- mentioned conditions as well as techniques for thermal protection structural elements testing for day-night and seasonal cycling.

Recommendations on carrying out the accelerated tests for climatic firmness are given. The approximate scheme of the main climatic factors affecting heat- protective material change in the experiment within the full-year cycle is presented.

It is noted, that experimental means for carrying out tests in the conditions of climatic influences must present a constituent part of the experimental means intended for the tests in the conditions of multiple-factor impact on of thermal protection materials.

The schematic diagram and photos of the test bench and its components used for heat-shielding reusable materials tests is provided.

The suggested methods and experimental facilities for conducting thermal-protective materials climatic tests on multivariable screen tests of tile-type thermal- protective structural elements can be used for consistent assessment of their working efficiency during ground tests. Ground tests of spacecraft units and plants can be conducted by simulating only the major external factors whereas secondary factors impact can be taken into account by introducing corresponding coefficients.

Based on the present-day maintenance conditions of modern missiles the paper reveals essential factors, which determine their reliability and readiness support at the required level. This, in its turn, allows establish that up to 70% of failures during instrumental control relate to guidance system, and more than half of them falls at the missile control system. The above mentioned problem manifested itself most acutely with short-range air-to-air airborne missiles. This implies the effective solution of missile control problem by employing guidance system functional control method and its realization based on hardware/software complex.

The paper suggests an original solution for short-range air-to-air airborne missile guidance-system loop technical condition, enhancing its functional control methodological efficiency, confirmed by methodological and hardware support synthesis.

Functional control scientific and methodological basics are determined by theory of similarity modeling and automatic control theory harmonic balance methods. The functional control effectiveness achieved with this method is determined by basic concepts inherent to the complete mathematical model structure, using the original inciting signal, generated by standard harmonic oscillations installation. These basic concepts include generation of such initial impact, which allow enhance missile guidance system controlled signals observability in system normal operation mode in space of parameters control.

The direct guidance system direct control time is one of the important parameters, related to its activation. This time is comparable to missile operation while intended application. It ensures the short-range air-to-air airborne missile specified life substantial saving.

Effectiveness of the methodological approach used by authors is supported by developing the guidance system software and hardware functional control complex that prevents introduction of changes to the guidance system hardware and sensors regular system. Thus, the possibility of practical implementation of the methodology, suggested by the authors, into field aerospace forces of the Russian Federation is guaranteed.

Employing analytical methods for computing spacecraft movement trajectories seems effective while solving a number of problems of practical importance. Analysis of the existing methods reveals that they are based mainly on mathematical models of spacecraft flight along a fore-and-aft plane, as well as some simplified spacecraft motion in space equations. It limits the possibility of their use while solving a number of space exploration problems of practical importance. The paper describes an analytical method for spacecraft atmosphere movement parameters computation. The scientific novelty of the developed method consists in transformation of a number functions in the form of recurrent piecewise-constant dependencies at the finite intervals of spacecraft flight trajectories.

After transformation of initial system of differential equations, we obtained the final computation dependencies for velocity and flight altitude, trajectory and course angles, longitudinal and cross range via the atmospheric density. Selection of such an argument, namely atmospheric density results from the fact that spacecraft flight situations can be identified based on calculations of this parameter with further recommendations for control decision-making. Based on the obtained equations we can compute not only the coordinates of spacecraft atmosphere movement, but evaluate the main characteristics, effecting design and technological decision making while a spacecraft design. Particularly, the fast evaluation of maximum overloads values, affecting a spacecraft in aerodynamic deceleration phases is provided. Analytic dependencies can be used while solving a number of variational problems in the conditions of preliminary definition of spacecraft control structure.

The tabular matter and graphical data are presented. Computation errors of spacecraft motion trajectory parameters are analyzed. It is shown that these computation errors do not exceed 2-3% with the total qualitative matching of obtained data and of differential equations numerical integration results. Employing of the developed analytical method allows obtain the highly precise computation results of spacecraft motion parameters in the atmosphere. The developed formulas provide high speed of calculations for a wide range of initial data, boundary conditions, and can form the base for spacecraft onboard control algorithms development.

This work relevance is stipulated by the necessity of airborne uncontrolled objet implementation accuracy to fulfill such tasks as forest fire extinguishing, large cargoes airlift delivery, etc. At present, conventional aiming systems do not provide uncontrolled object implementation effectiveness in full measure, since the onboard ballistic algorithm employs approximating equations and demonstrates low accuracy.

The authors suggest employ uncontrolled object motion complete ballistic model to improve onboard ballistic algorithm accuracy. The initial conditions can be obtained by determining uncontrolled object motion parameters based on radar signal reprocessing. These parameters determination can be realized with the algorithm, which description and structure are presented in this paper.

The paper presents computation results of the signal reflected from an uncontrolled object. These signals reveal that at the distances of up to 200 m secondary modulation harmonics of the first and second order are quite observable in the reflected signal spectrum, under condition of long-continued coherent integration of the signal.

The main advantage of this algorithm consists in the procedure of obtaining the unmanaged missile accurate initial conditions, based on the interpretation of the Dopplers effect together with complex application of known mathematical methods of signal processing. The reflected signal from uncontrolled object processing allows obtaining uncontrolled object launching (drop) angle, relative to the center of mass position, velocity and motion trajectory.

The paper presents a method for determination of overloads in the cross-section of long restricted articles, which can be employed at the initial stages of launching vehicles (LV) springing systems based on two-link tractor, while moving through rugged topography terrain, peculiar to Arctic zone of Russian Federation.

To evaluate overloads in in the cross-section of long restricted article the authors developed mathematical models of “two-link tractor — long restricted article” interdependent system, composed on the assumption of hitch mechanism infinite stiffness, when the LV presents infinite stiffness body, which practically eliminates the possibility of resonant speed modes occurrence while acceleration and moving with maximum speed.

The system of differential equations describing dynamic behavior of two-link tractor is divided into three less complicated systems of differential equations, which are solved by the original analytical method, namely combination method. This method is highly effective for dynamic systems study, if a differential equation does not exceed the sixth order. It presents an integral combination of symbolical and parameters variation methods. The symbolical method allows construct the resulting equation for the initial system of differential equations, and find dissipation and eigen frequency factors for the system under consideration. Parameters variation method, based on the solution obtained by symbolical method allows determine specific solution of the initial system of differential equations in the form convenient for the analysis.

The obtained results may be of interest to organizations involved in the design of viscoelastic suspensions not only for caterpillar tractors, but also for road and air transport, and exploring emerging overload of cargoes in extreme conditions.

According to the obtained results the conclusions on the expediency of operation of the hitch mechanism providing absolute rigidity of the coupling links of the tractor when moving on ground with periodic roughness in extreme operation conditions.

Two types of electric drives — electro-mechanical or electro-hydrostatic — are supposed to be implemented for оnboard systems of “more electric aircraft” with a great number of various kinds of electrical equipment [1] for controlling various functional elements [2]. The increasing number of implemented electrical equipment, electro-mechanical steering gear in particular, which phase currents cause electromagnetic interference (EMI) in the form of electric and magnetic fields with high-level intensities. The main source of radiated EMI caused by electric drives systems are power circuits bundles. Electro-mechanic steering gear power circuits' bundles consist of a pair of twisted or axial conductors with currents' of tens of amps [5, 6]. Correspondingly, they generate radiated EMI, which may cause signal distortion in aircraft onboard system communication lines and, hence, deteriorate its functionality.

The presented study is dedicated to of radiated EMI levels in the form of magnetic field harmonic components computation. These EMI are generated by phase currents in aircraft electro-mechanical actuator motor powering circuits, and crosstalk in the form of voltages in open conductors of double-wire communication lines.

 The presented spacing charts allow deduce that voltages and currents, which amplitudes are commensurable or even greater than valid signals values, occur in aircraft onboard cable system communication lines in the form of harmonic electric and/or magnetic field. The charts allow determining the safe distances between power circuits and open communication lines, wherein the levels of induced conducted interferences are significantly lower than information and control signals peak values in aircraft onboard system communication lines. It allows provide electromagnetic compatibility of high-power and low-power circuits.

The presented paper is a part of the research work on computation and simulation of electromagnetic interferences, caused by transients in aircraft steering gear system.

The рaрer suggests a control algorithm for voltage generation at induction motor windings by vector PWM. It reveals sрecifics of conventional vector PWM algorithm. Its is noted, that while through currents elimination with delay circuits a certain state occurs which allows identifying it as additional vector generation. The authors suggest a control algorithm with extra vectors generation as a through currents elimination technique. The рaрer comрares the suggested technique with conventional, and demonstrates that the develoрed algorithm using extra vectors allowed eliminate through currents of a first genus, decrease amрlitudes of high-order harmonics, and ensure рhase and рhase-to-рhase voltages рarameters similar to the conventional technique. Simulation of the suggested technique was carried out, and its results revealed that рhase current sine waveform could be ensured not only by increasing the number of generated vectors in one sector, but also by introducing extra vectors.

Conventional and suggested techniques reveal that рhase and рhase-to-рhase voltages characteristics as well as рhase current are similar, but the number of high- order harmonics is less than with conventional one. The breadboard tests revealed that the develoрed algorithm did not lead to shaft whiррing. Inverter inрut current herewith is less relative to the conventional vector PWM technique.

With on state of intermediate vectors significantly less than on state of base vectors the рossibility to attain рositive features рeculiar to the conventional technique, but eliminate a number of its drawbacks. The suggested technique, in рarticular, allows eliminate through currents, and gives more рossibilities of vector PWM imрlementation. Extra vectors on state duration control, rather than increasing the number of generated voltage vectors, allows ensure рhase current shaрe more close to sinusoidal.

Creating a competitive technology requires implementation of new materials with high specific mechanical properties. Conventionally, such materials are produced by introducing alloying elements, which form strengthening phases within the base metal structure. This approach usually results in the mass gain, because the hardening phase density is often higher than that of the base material. The mass of material can be reduced by introducing it into the volume of structural defects, such as pores. Due to high damping properties, low thermal conductivity, high sound-insulating ability and good moisture resistance, the porous materials are widely used in industry [1-7]. With existing porous aluminum, manufacturing technologies its strength properties decline takes place. However, with porous structure ordering the strength properties of finished products improve [8-10].

Thus, the goal of the presented work consists in improving specific mechanical properties, yield strength in particular, of the material by introducing orderly arranged pores.

This study employed deformation processes finite element modeling with engineering analysis pack MSC.Marc to determine an optimal porous structure [11-12].

The study of porosity and a type of porous structure effect on mechanical properties was carried out with the following types of porous structures: square, field-interleaved, square with a pore in its center, triangle and hexagonal.

With porosity of 0.4 to 0.5% porous samples FEM yield strength matching with compact material FEM samples yield strength is observed. With further porosity decrease growth of yield strength is observed for all types of porous structures. Maximum yield strength increase of 1 to 2% was achieved with porosity of 0.1%.

The blanks for all the samples were cut from the aluminum alloy A5 sheet using laser cutting complex. All the obtained blanks were decollateв into three parts. The first part was left intact as a compact material sample. In the second part of the blanks, the ordered porous structure was obtained by laser burning. In the remaining samples, the porous structure was obtained with CNC milling and engraving machines with the drill diameter of 300 µm.

The finite modelling and real uniaxial tensile tests results matching is observed.

Nowadays “AMg6”, “D16” and “AMn” aluminum alloys are traditionally used for space technology. Application of new advanced aluminum-based materials with of rare-earths additives instead of traditionally used alloys would enhance the electronic components protection from the space ionizing radiation due to alloying with high radiation absorbing elements. Whereas chemical composition manufacturing technique optimization will improve, alloys' mechanical properties compared to conventionally used, which will allow decrease weight and size parameters of the design.

Tests carried out by Russia's space industry leading organizations revealed significant preeminence of new alloys compared to conventionally used with regard to protection against outer space ionizing radiation properties, and corresponding to them ability to chemical electroplating. Aluminum based alloys specific mass with rare-earth additives is 2.9 g/cm³ on the average.

This paper is focused on the study of the three different alloying systems: 1 – Al-Dy-La-Cr-Zr, 2 – Al- Ce-Cr-Zr, 3 – Al-Mg-Sc-Zr-La-Ce; with rare-earths content not exceeding 11%, 7% and 9% by weight respectively. Each of the studied alloys, regarded as a material for spacecraft electronic equipment casing has a number of advantages and disadvantages. Increasing the rare-earth metals content in the alloy we can attain both better protective characteristics against space ionizing radiation, and aluminum based alloys with rare-earth additives welding properties improvement. Tough their density herewith will increase. Thus, it is necessary to pay special attention to improve mechanical properties of the basic metal and welding joints to prevent weight and size parameters of the design. Mechanical properties improvement with density reduction may, in some degree, be achieved by rare-earth aluminide phases' dispergating and increasing their density distribution in the alloy groundmass.

The authors have developed a method for computing a number of intermetallic phases (T₁ and δ′- phase) of Al-Cu-Li-Mg system alloys based on measuring α-solid solution lattice periods; Vegard’s law, linking up solid solution lattice period value with alloying ingredients content in it, as well as chemical and phase content equations. Lithium content in solid solution serves herewith as variable parameter. Quaternary Al-Cu-Li-Mg system alloys quantitative phase analysis method is based on the assumption that all magnesium resides in the solid solution. This fact is considered by introducing the relevant term into equation for calculating the solid solution lattice period. This is the only difference fr om the previously developed similar method for ternary alloys Al-Cu-Li. The paper shows that the developed method can be effectively used for quantitative interpretation of thermal and thermomechanical processing impact on alloys’ phase content study results, as well as while Al-Cu-Li-Mg system alloys content optimization. This method allowed us to compute the relation between periods of solid solution lattice and the amount of intermetallic phases for 29 Russian and foreign industrial alloys of various generations. The paper reveals the existence of linear dependence of relative quantity of intermetallic phases in alloys  from the atomic concentrations of lithium and copper (magnesium)  in these alloys. It shows also, that relation between δ′-phase and ternary phases is determined by the atomic concentration of lithium and copper. The authors suggested new Al-Cu-Li-Mg — alloys classification, wh ere all alloys should be divided into five groups, differing from each other by the double δ’-phase and ternary phase shares , or  ratio.

According to this classification, all the alloys are divided into five groups. The first group includes Al-Mg-Li alloys, for which the phases ratio . For the second group the ratio  varies from 2 to 3; for the 3rd group — from 5 to 7; for the fourth group — from 7 to 8, and the fifth group — from 11 to 17.

The goal of the paper consists in titanium alloys semiproducts macrostructure quality evaluation technique improvement.

Active standard ten-point scale of macrostructures was developed based on – grains of strictly equiaxial shape specific to strain-free state of the alloy of sheet-like intragrain structure.

It is well know macrostructure we can see only those structure elements, which size exceeds 100 –150 micron (i. e. the ones exceeding the eye resolution capability).

Macro- and microstructure evaluation of a large number of serial semiproducts and laboratory samples revealed that not only – grains could be visible on a microstructure, but – colonies as well. It was established, that while checking a established, that the shape and size of the grains in the observed macrostructure depended on – grains and  – colonies in the microstructure.

Direct dependence of a macrostructure character from its microstructure was revealed. The paper shows that macro grain size and its tonality (degree of brilliance) depend directly on parameters of the microstructure, forming while deformation and heat treatment processes at temperatures of -or + – area. Correlation between the grain maximum longitudinal and diametrical sizes (the degree of non-equiaxiality – K) is clearly associated with physical degree of its deformation. This is another important parameter of macrostructure evaluation besides the grain size itself.

By deformation, the macro grain tonality or its degree of brilliance changes together with the macro grain shape. Interrelation between degree of brilliance of a macrostructure under study and with its microstructure was established.

The author suggests classify a macrostructure according to its tonality (degree of brilliance) by four types:

• Absolute brilliant – a typical macrostructure peculiar to allows with recrystallized or slightly malformed – grains which size exceeds 100 microns.

• Brilliant with fog elements – observed in alloys with medium degree of deformation (10-35%) in +  – area. Within on – grain one can observe micro areas withvarious degree of – phase spheroidizing development (from globular to practically non- spheroidized, plate-like shapes of the particles).

• Fog with brilliance elements – peculiar to the alloys malformed in +  – area to the degree of 40-55%. For the most part this structure is globular or globular plate-like. In some locations it preserves oriented character of – phase excreta, which in case of their large size are responsible for appearance of these brilliant locations in the macrostructure.

• Absolute fog-corresponds to globular or globular plate-like microstructure.

To improve titanium alloys evaluation objectivity and unambiguity the author suggest introduce quantitative estimation based on three parameters, namely grain size, and the degree of its non-equiaxiality and tonality.

The next step to titanium alloys production quality improvement consists in working out requirements to macrostructure based on quantitative estimation of its parameters.

The paper suggests methodological approach to fundamental and research priorities portfolio forming. Such an approach is based on expert selection procedure. The authors formulated thematically oriented verbal and numerical scales for qualitative selection criteria groups. This criteria grouping is organized according to thematically similar features, significant while new perspective aircraft systems design.

Due to the complexity of solved selection tasks it is reasonable to implement quality criteria system, described by 3-5 criteria groups with 3 to 7 criteria in each group. It allows convenient, transparent and comprehensive presentation of information to the expert in necessary and sufficient scale.

Group-1 represents usefulness and importance; Group-2 represents resource intensiveness and resourcing, and Group-3 represents stability and manageability.

With allowance for expertize complexity and supposed relative inconformity of experts opinions when evaluating significance of particular researches trends, we suggest selection procedure based on the Ansoff's theory of weak signals. Here, with allowable level of experts' nonconforming opinions, the individual opinions of competent experts with high estimate of particular FERs are taken into account. Here, core index (CI) and concordance index (DI) serve for the generalized selection measure. In this case we suggest FERs grouping in the following way.

FER-1 are the trends with experts' high estimation by CI with high DI value. FER-2 are the trends with relatively high estimation by CI with relatively low DI values. FER-3 are the trends which CI is better than this for FER-2, with DI lower relative to FER-1 and comparable to this for FER-1. FER-4 represents such FER trends, which received consensus on lower importance, either special opinion was expressed by experts with relatively low authority.

The authors suggested to form the portfolio not only by the trends with high CI and DI values, but consider FAR-2 trends (with priorities higher than this of FAR-3) as well. This approach allows us to identify and select among the priority research areas with potentially high efficiency, albeit with relatively high level of risk. The proposed approach also makes it possible to make informed decisions in a limited time based on authoritative (respectable) peer review. The method is oriented for use in decision support system.

The paper presents a technique for a number of work places optimization at the enterprise with variable product release program. The developed technique is based on simulation and experimental design. The paper considers the operation of the enterprise manufacturing several kinds of products by assembling either purchased components or produced at this enterprise. The simulation model developed in the course of this study allows build and optimize manufacturing resources under various variants of enterprise's target figures.

The model was built with AnyLogic program, which allows specify time intervals at every stage of manufacturing either major product, or associated items. There is a possibility to model the situation with various number of assembling departments.

Based on the built model the authors carried out the optimization experiment, which allows compute an optimal number of equipment for the specified work-order quantity for all types of products. The paper suggests goal functions with productivity optimization. Using this instrument the results for each experiment were obtained by varying values of run-out production plan. It is found on what production volumes minimum quantity of equipment is optimal, and at what moment the number of working places should be increased. It is also determined that maximum possible quantity of equipment under specified production volume boundaries is not necessary.

The obtained results were processed according to the experimental design technique. The equation for corrected production effect computation as function of a number of assembling departments and products production volume. The proposed method is universal and can be applied for various types of production. The developed technique can be used as one of the instruments while developing the system of managerial decision making.

At present, creation and implementation of a system of indicators for monitoring R&D processes is necessary, since the degree of highly efficient estimation of economic security of an enterprise, and formation of economic mechanism implementing the complex of necessary measures on prediction and preventing a danger, corresponding to the scale and threat environment to aircraft industry in the aggregate.

The main objective of the innovative activities economic security level consists in timely analysis and monitoring of a complex indicators system, inclusive aircraft industry specifics.

Development and implementation of economic mechanism for innovative activities economic security provision in aircraft industry will allow reveal: insufficient certainty of a forecast at various R&D fulfillment at stages; excessively enlarged and averaged character of labor intensity rate and expenditures, new objects' operational service norms, reliability and durability; insufficient comparability of new objects with selected prototype objects, or the lack of scientifically substantiated techniques for this comparability economic evaluation while effectiveness indicators computation; the lack of exact information on all spheres of R&D results supposed implementation and their scope of their implementation; the difficulty of extracting the share of economic effect related to this particular technical solution, the specified object, used as a constituent part of a more complex technical system.

A corporate governance system developing method is offered within the framework of this paper.

This newly developed method aims at estimating the corporate development level of a space-rocket industry enterprise, as one of the crisis-proof enterprise governance tools.

American and German corporate governance models are distinguished as the most popular models. Russian corporate governance practices are equally distant ideologically from both.

To analyze the corporate governance system at the rocket-space industry enterprise the authors developed their own estimation method, consisting in accounting for a number of specifics of integrated structures and individual enterprises functioning.

Basic methodological approaches to optimal assessment method of corporate governance system development at the enterprises and integrated structures of rocket-space industry are considered.

The developed method account for specifics of the companies and allows carry out their comparative assessment in conditions of differences in structure, scope and lines of activity.

This method represents a multilevel assessment system by two blocks — the level of corporate governance implementation and the level of the corporate governance system formalization.

In accordance with the results of the study, the authors obtained the optimized branch method of estimation of the level of corporate governance system development in the interests of crisis-proof management.

This paper is dedicated to the study of economic relations associated with socio-economic inequalities and its component — financial inequality, which is reflected at the global, state and regional levels, as well as the levels of enterprises, organizations, institutions and individuals.

In modern conditions financial inequality worsened dramatically with the development of the world economy, deepening of global economic relations; development of financial relations; strategic amalgamations between corporations, alliances, unions; deepening of contradictions between rivaling states; globalization; labor mobility, goods and services; formation of a unified information system.

The goal of the paper consists in emphasizing the importance of financial inequality as an economic phenomenon, analyzing its dynamics, developing recommendations on its mitigation and implementation of State regulation by using aggregate financial instruments, such as budget, taxes, public debt, etc.

While carrying out the studies and tis paper formation the author employed scientific methods of analysis and synthesis of theoretical views and concepts, economic and statistical methods to assess the impact of various factors on the financial inequality, as well as scientific heritage on this issue. The study was conducted based on the principle of historicism.

Economic theory principles, scientific concepts and theoretical evidence of national and foreign scientists served as this work methodological basis.

The results of the work consist in generalization scientific ideas and viewpoints of national and foreign scientists and economists, and determining socio-economic inequalities as inequality of wealth associated with the degree of realization of human rights, democracy, free market and effective State laws.

Theoretical research, opinions and viewpoints contained in learned writings on the issue of financial inequality will supplement the content of academoic disciplines related to “Finance and Credit”, taught to students of higher professional education.

The results of the study, based on analysis of financial statistics on income, inflation, taxes, debts, are rather efficient at the macro-level while forming the state policy.

The results obtained in the course of the study are aimed at State regulation of financial inequity mitigation through implementation of fiscal measures to the social protection of the population. The state support during economic crises and the need for economic regeneration will affect significantly the financial inequality mitigation.

The paper suggests creative destruction of informal relations dominance, transparency of state accountability, formation of a balance between State regulatory effects on financial inequality and self-regulation on the part of businesses and individuals.

The main issue of this article consists in analyzing the main electric network emergency states and to designing a protective system model, which is able to minimize or fully avoid their aftermath.

The results of failures analysis at various power plants and installations allowed separate out the basic types of emergency operating modes. They include sudden voltage dropouts, voltage waveform fluctuations (flickering), rolling blackouts and presence of significant reactive power abundance in a power grid. The rolling blackout presents the greatest danger due to its aftermath. The analysis of emergency modes occurrence revealed that most commonly they arise due to insignificant event, leading to avalanche-type emergency growth. This fact is reflected in the presented algorithm. Moreover, most commonly, these emergencies can be eliminated with the timely reaction of the personnel. However, as the practice indicates, these specified nonsignificant factors were ignored by maintenance staff.

Two-level of consumers' complex protection model for emergency elimination and its aftermath mitigation is suggested. Both parts are autonomous and can be set separately, or in conjunction. The first part of the system is responsible for the reactive power compensation in the power grid. It differs from the existing prototypes by smaller size, cost and asymmetric structure for reactive power compensation in wide range. The paper presents voltage and power balance graphs at the object before and after compensation. The presented data proves that implementation of such installations allows reduce rolling blackout occurrence probability.

The second part of the system represents from rolling blackout protection controller, which, in case of any power grid section overload, or voltage dropout, analyses the states of consumers and turns off those of lower priority. This helps avoiding entire system cascaded failure occurrence.

The presented system both as a whole and in separate parts presents interest for industrial electric energy consumers from the viewpoint of spoilage minimization occurring due to power grid failures.

This work aims at developing a dynamic object multidisciplinary optimization, namely maneuverable aircraft horizontal empennage.

Horizontal empennage is a complex technical system with a great number of parameters described by the equations belonging to various scientific disciplines. Thus, the developed method represents a multidisciplinary optimization method. The efficiency of the horizontal empennage can be evaluated by the value of the moment produced by the horizontal empennage and the pitch moment gradient magnitude, which can ensure the aircraft balancing and performing a specified maneuver. Parameters necessary for optimization in the framework of the specified problem were determined and their ranking was performed by weight factor determination for each parameter. Then the goal function for horizontal empennage parameters optimization was created. Various approaches to supersonic aircraft horizontal empennage parameters optimization, such as method using Pareto principle, or method using Hurwitz criterion, were studied and realized. Analysis of operation of the above mentioned optimization methods in the context of the specified problem revealed their insufficient efficiency. With the aim of improving the obtained results a new optimization method was developed and suggested. This method employs the valuation of several valuation functions to obtain optimal solution. The effectiveness of the developed method is demonstrated using various input data sets, and the effect of various weight factors parameters on obtained result was studied. Its operation results in horizontal empennage optimal geometric parameters, formed automatically with CAD system Siemens NX 7.5 “Modeling”.

The paper considers computational results of moments relative to main rotor axial hinge caused by action of inertial and aerodynamic forces occurring on the blades. The computational method is based on the generalized disk vortex theory of rotor in an oblique flow in its most simple version, when the air load over rotor disk is assumed constant in computing inductive speed vector components. Flapping movement coefficients of hinge-suspended in thrust plane rotor blades are defined by analytical relations with an accuracy to the first harmonic of Fourier series. This assumption reduces the problem of hinge moments calculation with specified rotor control law to the method of successive iteration on thrust force ratio.

As an example, the authors considered the rotor blades operation of helicopters Mi-34 and Mi-28. They studied the change of hinge moment value depending on the blade azimuthal position, and evaluated each components contribution to it.

The wobble plate rod strain dependence on flight speed, blade angles and control action was analyzed. It was established that with wobble plate ring deviation in forward longitudinal direction law of azimuthal strain variation demonstrates harmonic character with pronounced maximum near the retreating blade. With wobble plate ring deviation in transverse direction the similar dependence is of the same character, but its amplitude is negligible compared to longitudinal control.

Calculations were executed for helicopter main rotors with various structures of a hub and blades. The results were compared with experiments and calculations of the other authors. It is shown that mathematical model of absolutely rigid blade combined with disk rotor vortex model allows evaluate control system loading level at various helicopter flight modes with adequate accuracy. The ways of model improvement allowing define the obtained results more exactly are outlined.

The article depicts methodology and gives recommendations on complex geometry air intake modelling process on the example of uncontrolled air intake. These recommendations have been elaborated considering many requirements and suggest the procedure, which could be used during air intake channels integration into aircraft layout.

Main requirements to the geometry (equivalent angle of subsonic diffuser) and relative length (Lch/Dent, Lch — channel length, Dent — diameter of engine entrance), as well as parameters of air intake entrance channel are imposed after all boundary conditions, such as air intake input square (F₀), air intake throat square (F_th) and optimal graphics of air intake duct squares are set. The suggested method for deceleration system integration allows the transfer to intake duct surface constructing with minimum iterations.

The practical part of this work includes recommendations to construct optimal complex duct surfaces using “Unigraphics” with allowance for their maximum optimization. We suggest to use “dynamic changes” method while designing air intake, i. e. changing associative construction with construction parameters changing (shape-generating parameters). Based on created geometry we plot the graph of air intake duct squares according to the squares obtained duct cross sections. It allows demonstrate the dissimilarity of actual duct geometric parameters from the presumed (optimal).

The above-described procedure allows select air intake and air duct location of supersonic jet with allowance for a large variety of layout limitations and gas dynamic requirements. It allows integrate air intake ducts of complex space form into aircraft layout under development with minimal time consumption. The suggested procedure of “dynamic” air intake duct geometry changing allows control changes of its geometric characteristics (graph of air intake duct squares function) and select thereby the duct parameters with subsequent optimization.

The important goal of GTE combustion chambers and installations design and workout consists in provision of the specified durability and reliability. In response to this problem, the temperature of the combustion chamber major elements should stay within the operating temperature range of their structural materials, and their deformation should not exceed the specified values.

Combustion chamber flame tube walls are some of the most heat-loaded elements. Thus, the problem associated with the study of their stress-and-strain state is particularly up-to-date. It is worsen by the existing tendencies aimed at increasing compression ratio of the compressor and gas temperature at the turbine input. Temperature distribution over compressor flame tube surface determining in the course of complete product bench testing presents a complex task. The effective method of rectifying the above said problem during gas turbine units consists in numerical modeling methods implementation, which requires developing procedures of their implementation.

Such procedure was developed while GTD combustion chamber for terrestrial surface application design. For its realization we used geometrically conjugate 3-D model of the combustion chamber, including both air-gas channel necessary for gas dynamic processes modeling, and flame tube walls with multilayer thermal-protective coating for heat transfer computation. Combustion chamber operating procedure mathematical model was developed earlier and passed validation process. Simulation was carried out in ANSYS.

Temperature distribution on the flame tube wall was obtained by computation. Based on the analysis of the obtained results we managed to reduce maximum flame tube wall temperature to the required value at the expense of apertures areas redistribution between cooling system strips. Stresses and deformations occurring due to flame tube walls heating were determined as well. It was revealed that maximum stress occur at cooling apertures locations. The value of calculated strength factor equals 2.7.

The developed procedure for combustion chamber flame tube walls thermal-and-stress states determination can be implemented for various combustion chamber desings, materials and multi-layer walls. This procedure allows predicting the most dangerous temperature zones on the flame tube walls and burn-out in this zones prior to bench testing.

The purpose of this work consists in clarification and generalization of two-shaft gas turbine engine rolling bearings vibration diagnostics method. The paper gives the review of the well-known problems associated with gas turbine engine rolling bearings vibration diagnostics, and considers methods employed while carrying out diagnostics. We sel ected the procedure using hand rolling of one of the engines' rotors for our studies. The paper considers possible problems occurring during diagnostics. For the studies we assume the same frequency range that is used while hand rolling. The vibration signal measured prior to the complete stop of the engine is generated under identical conditions. Such conditions are characterized by very low forces, causing forced oscillations. Thus, we can neglect these oscillations. The signal for the study carrying out was obtained fr om the engine with damaged bearing. The degree of the damage development was rather high. The rotor decrease in speed, and short time interval do not allow employ conventional rolling bearings spectral vibration diagnostics methods. Analysis of the signal temporal realization reveals the absence of considerable shock load. Spectral analysis reveals the presence of only one significant harmonic in coincidence with the engines natural frequency. this natural frequency is characterized by the high value of deformation latent energy accumulated in in the bearing. The absence of the shock load points out that natural oscillations excitation does not occur. The inference is drawn on the presence of self-sustained oscillations in rotor rundown mode, developed at natural frequency. As long as the oscillating processes in the engine are almost finished, the natural oscillations cannot obtain energy for the amplitude rise. In view of natural oscillations amplitude smallness their impact on bearing degradation is insignificant. Self-sustained oscillations process phase diagrams are plotted. Strap brake is a literary analogue of the being obtained self-sustained oscillation system. Despite the vibration signal non-stationarity, spectral analysis allows obtain reliable diagnostic results. The conclusion has been drawn about the possibility of applying the considered phenomenon for vibration diagnostics of roller bearing dual-shaft gas turbine engines.

The article considers the development of small gas turbine engine combustion chamber draft design. Though it occupies relatively short period of time, it is an extremely important element in reference to the engine lifecycle period. The draft design results allow obtain all necessary information about both the combustion chamber as a whole and its constituent parts.

The small-sized gas turbine engine combustion chamber draft design allows determine its shape and the features of its embodiment. Geometry selection and integrated parameters evaluation with account for turbocharger components composite action is important due to incomplete information on these matters.

The paper presents the results of the statistical data processing of geometric parameters and integral characteristics of the GTE combustion chambers. Complex correlation functions were obtained by methods of mathematical statistics. Correlation functions describe the structure and geometry: primary zone operating parameters and combustion chamber integral characteristics such as: flame tube volume (VFT), relative length of flame tube(LFT/VFT), nozzles relative pitch, combustion efficiency, residence time, forcing coefficient (K_V), thermal factor (Q_VP) and relative areas of internal and external channels. Changing of the abovementioned values depends on consumable complex (G_CC). Table below presents results of the statistical data processing. Most coefficient of determination values lay within the range of 0.06 to 0.7.

Not all of the dependencies have high coefficients of determination. However, this does not exclude the possibility of their use in the preliminary assessment of the gas turbine engine combustion chambers structural and integral parameters.

Solar thermal propulsion (STP) is considered as means of inter-orbital transportation from low Earth orbit into geostationary orbit. The payload insertion time for conventional STP usually equals approximately 60 days. STP contains high-temperature concentrator-absorber system (CAS) as a power source, with possibility of using multi-staged absorber of non-isothermal type with higher optical-power efficiency. Ballistic efficiency of the upper stage with STP grows with the increase of extent of CAS not-isothermal properties and can exceed 1.5-2 times the efficiency of liquid propellant rocket engines (LRE).

Ballistic efficiency of upper stage with the STP is determined by relevant parameters of the CAS, to which we can assign concentrator accuracy parameter and hydrogen heating temperature in CAS, as well as permissive conditions of CAS's sun orientation. Selection of CAS expedient parameters can be realized, in some cases, different from those optimal with allowance for technological limitations under permissible reduction of ballistic efficiency of solar upper stage compared to LRE implementation case.

Considering that transportation system, employing combination of high and low thrust engines with adequate insertion time of 60-120 days may compete with STP as inter-orbital transportation system, it is expedient to estimate the ballistic efficiency of solar upper stage at lower flight time. The problem simulation demonstrates the capability of reducing the payload injection time to 20...40 days at high ballistic efficiency of solar upper stage (in case of extreme non-isothermal multi-staged CAS, the payload mass is twice higher compared to LRE). STP optimal relevant parameters herewith change towards CAS simplification. The CAS Sun tracking conditions also become simpler.

STP optimal parameter values for various time intervals of inter-orbital transfer are presented. The possibility of providing high ballistic efficiency at the flight time of 20-40 days with STP implementation with non-isothermal multi-staged CAS, compared to the transportation system employing combination of engines of large and low thrust, is shown.

The paper considers parachute system with pivotally suspended load movement in vertical plane. When deriving the equation of motion the parachute system is presented as a mechanical system consisting of two solid bodies connected by ideal linkage joint (deadeye), i. e. parachute and load. Parachute system free oscillations analysis is carried out with the equations of the simplified disturbed motion model, where translational movement with constant speed and constant parachute canopy and load incidence angles.

As partial movement mathematical model the equations of first approximation system are accepted, and the sought non-linear model is obtained by substituting in these equations normal components of parachute canopy and load aerodynamic forces coefficients linear dependencies from their incident angles by non-linear dependencies. The dependency of parachute canopy aerodynamic force normal component from its incidence angle herewith takes the form characteristic for the case when canopy fabric has low permeability. After harmonic linearization of these non-linear dependencies the thus obtained system of non-linear equations for simplified non-linear model falls into two systems corresponding to constant and variable components of the sought solution. Equations linking amplitudes and center displacements of the sought oscillations are derived from the equations containing constant components, while equations linking amplitudes with oscillation frequencies and inequalities, determining stability conditions of the sought oscillations are derived from the equations containing variable components. Implementation of the suggested procedure of parachute system with pivotally suspended load with allowance for aerodynamic load is illustrated by numerical example. The boundaries of canopy incidence angle initial values that provide damping oscillations mode of the considered parachute system near the specified unperturbed motion are determined with or without considering the load aerodynamics. To evaluate the results of theoretical calculations the paper presents the results of numeric integration of the initial equations of the parachute system movement.

The paper considers the current state-of-the-art of uncontrolled axisymmetric object throwing together ballistic support, and analysis of the existing techniques for their trajectory calculation when launched from an aerial vehicle (AV). The main parameter specifying uncontrolled object motion relative to the center of mass is nutation angle δ , which initial value can achieve the threshold values (abou δ₀= 30°) and depends on combinations of its projection, such as angle of throwing λ₀, separation velocity, environment disturbance, launcher vibration etc.

To improve accuracy of the problem solution aboard an AV, we suggested to use full ballistic motion model during its separation from the carrier, and the initial conditions to solve it could be obtained according to the ballistic characteristic determination technique of uncontrolled objects being offered. This technique can be implemented under condition of availability of testing system, which make-up and structure are presented in the paper.

Main advantage of the testing system is implementation of proximity sensors comprising a set of sectors realized in the form of perpendicular arranged rows of photo-detectors and radiating elements.

The paper presents for uncontrolled object motion trajectory with initial nutation angle computation results, which show that when throwing 30-millimeter uncontrolled axisymmetric objects rotating about its axis with initial nutation angle values not-to-exceed 10° to a range of 1500-2000 m, the range spread up to 170 m and lateral deviation spread up to 20 m are observed.

In range environment, the proposed technique for ballistic characteristic determination of uncontrolled objects enables to obtain dependences of initial uncontrolled object nutation angles versus different conditions of their launch. The resulting experimental dependences can be utilized to obtain accurate initial conditions required when integrating differential equations of full ballistic motion model for uncontrolled objects.

The paper discusses the issues of pilot's dynamics detection and its accounting in the process of glide-path capture accuracy appraisal, as well as accounting pilots dynamics effect on aircraft precision of movement while glide path movement and flaring-out. We suggest solving these issues based on complex mathematical modeling.

The pilot’s behavior based on representing the pilot as an optimal non-linear regulator, and experimental data search of generalized criterion of pilot’s control activities were considered in details. The obtained formal criterion with derived weight factors enables realization in ACS the algorithm identical to pilot’s control activities while aircraft piloting in the form of a direct problem of getting from an arbitrary point in airdrome area on approach glide path.

Three types of stages of landing and corresponding pilot’s models, such as glide path capture, movement on the glide path, and flaring-out before runway touchdown. Modeling and algorithmic analysis of various aircraft thrust control laws allowed selecting the most expedient altitude of flaring-out starting on the assumption of flaring-out altitude valuation permissible error maximizing principle. The obtained permissible areas of initial flaring-out altitudes provide maximum pilot’s comfort in case of forced manual landing mode transition.

The aforesaid solutions enable desired trajectories shaping and algorithms realizing automatic landing according to anthropocentric principle, providing, if necessary, fast transition to manual control mode in case of automatic control rejection, adapting to the current situation.

Anthropocentric approach allows the pilot to operate under minimum psychological tension, since while automatic control he observes the movement, which he would realize himself in case of the necessity of manual mode transition.

The article analyses the state-of-the-art of aviation guided missiles development. Today guided missiles guidance systems thermal imaging coordinators have widespread application. Reading units with digital image processing and fiber optics gain maximum acceptance.

The problem of guided missile neutralizing at a safe distance from the defended aircraft was solved. An approach allowing increase the safety of a defended aircraft through ballistic support of active protecting elements casting was considered.

This method employs the automatic ejection unit allowing ejection of active protecting elements in the required direction. Furthermore, the on-board information systems, computer, mother ship measurement sensors and ejecting unit are integrated into unified complex.

A model of active protecting element movement for on-board ballistic algorithms synthesis was created. It allows develop mathematical tools defending complex dataware while active protecting elements implementation, scientific and methodological tools for their effectiveness evaluation. Rational implementation of active protecting elements as a part of aircraft system was substantiated. The developed algorithm allows realize the conditions of active protecting element implementation in real technical applications, operating within the aviation system.

While determining the parameters of guided missile movement relative to an aircraft, the trajectory of the rocket movement was predicted according to the guidance mode employed by the missile. For this purpose the guidance system obtains the information on coordinates and other aircraft and missile movement patterns at every time instant. It sets the character of their interrelation, determines the degree of this interrelation disruption. Based on this information it forms the parameters and control signals, providing the required movement of a rocket to an aircraft.

Using navigation and weapon-aiming system computing capabilities, together with on-board defensive systems the time of a missile closing-in with and aircraft was computed. Ballistic model of active protecting element casting ensuring its encounter with a guided missile was developed.

The result of the algorithm operation provides neutralizing of a rocket in the air, thus preserving an aircraft performance ability.

The paper presents the bundled software meant four wheeled off-road vehicle convoy itinerary selection, while unmanned aerial system (UAS) re-deployment from the deployment site to operating zone. The main feature of this bundled software consists in the possibility of its use not only for road travel, but also in the possibility for travel the part of the route on a country road and off-road.

The performed analysis revealed that it is expedient to select easily realized algorithm as the theoretical basis, such as Dejkstra algoritmh, as well as Floyd-Warshall algorithm and Bellman-Ford algorithm.

UML models of the system under development were constructed.

The paper presents the test example of forming characteristicsof transport means constituting the convoy. It allowed determine for each road layout sector its negotiability by the convoy and the convoy travel time along the sector, using the built-in calculator. The authors used simplified computation algorithm to calculate route sectors negotiability parameters and the speed (time) of their passing those sectors. Then the negotiability of a certain sector was determined. Finally, a graph of roads negotiability was plotted. The shortest distance between the UAS locationoriginand redeployment destination point is searched on the plotted graph.

The developed automated information system allows:

— Reduce transport means movement constituting a convoy moving off-road route planning labor intensity due to automation of the majority of operator's functions;

— Increase the quality and accuracy of calculations, needed for road layout graphanalysis;

— Reduce the personnel, necessary for execution of work on UAS convoy movement planning, costs;

— Obtain up-to-dateaccountingon variousUAS movement route measures of efficiency in real time mode to take managerial solutions.

The study of helicopter dynamic stability is associated in most cases with such a phenomenon as ground resonance, which presents helicopter self-oscillations with increasing amplitude. The origin of this swaying consists in interaction of blades oscillations relative to the hinges, vertical in particular, with helicopter body on chassis oscillations.

Since helicopter presents a complex mechanism, assumptions and simplifications allowing saving computing resources and with sufficient accuracy easily correlated with the experiment were introduced while mathematical models building. At present, a researcher possesses considerable computing resources. Thus, one can afford building much more complicated models, allowing solving the problem of «ground» resonance computation in more detail.

The paper presents helicopter mathematical model, which body has six degrees of freedom, and flexible blades with three degrees of freedom in the attachment point to the hub. Helicopter alighting gear (chassis) is presented in the model by flexibility matrix. Helicopter equations of motion were obtained using second order Lagrange equation, and blades flexural oscillations equations were obtained with widely known Galerkins method. Flexible blade mathematical model considers only the first three forms of hinges flexure-flexure-torsion oscillations.

Following the above-described mathematical model the complex of programs was developed using Maple and MATLAB. Within the range from zero to main rotor operating speed computation of helicopter dynamic instability zones on the ground was made. Comparison of the results obtained by R. Coleman method and mathematical model with rigid blades for ANSAT helicopter revealed sufficient convergence. Mathematical models with rigid and flexible blades developed by the authors allow determine additional instability zones.

The model with flexible blades allows revealing a number of additional instability zones, which may have great significance.

Heavy cargo parachuting off an aircraft is the fastest wat to deliver it to the uttermost remote and hard-to-reach areas. However, the pilot faces difficulties steering an aircraft while discharging a heavy cargo, due to the aircraft center of mass considerable changes. This problem aggravates now for several reasons. Firstly, the weight increase of heavy cargo, armored vehicles in particular. Secondly, the increase of speed range and dumping heights. The possibility of an “air start”, i. e. a spacecraft airplane launching at high altitudes, as well as rescue vessels air drop is considered. A free drop from deck levels of 5 ± 2 m is rather difficult due to the earth immediate proximity.

The possible solution to the piloting complexity problem consists in automation of this process. It is necessary to justify scientifically and methodologically the need and possibility of an aircraft control automation while heavy cargo extraction. Thus, the object of the study is the process of aircraft control while a heavy cargo extraction, and the subject of the study is its automation.

The research task are as follows:

1. Revealing the gist of the problem of the MTA flight parameters violation of operating limitations while heavy cargo extraction.

2. Analysis of the factors affecting the MTA piloting while airlift delivery.

3. Сonsidering the existing ways of solving the problem of the MTA flight parameters violation of operating limitations while heavy cargo extraction.

4. Analysis of the possible ways of MTA automation while airlift delivery.

The need for flight control automation was proved based on the study of the MTA flight parameters violation of operating limitations while heavy cargo extraction and consideration of the factors affecting this process. The results of the analysis of the existing ways of the abovementioned problem solution and possible ways of flight control automation allowed develop adaptive control algorithms for aircraft control while heavy cargo extraction.

It must be emphasized, that the detailed analysis of the abovementioned problem and rational ways of its solving under modern conditions were made for the first time and, furthermore, it seems expedient to carry out further studies on developing the automation technique based on adaptive algorithms with current identification.

Sufficient outside world view fr om pilot’s flight station is one of the essential conditions of the safe and comfortable aircraft piloting, including a helicopter.

Horizontal and vertical look-up angles from the basic directional point C explicit numerical values are specified by regulatory documents, and by FAR-29 Aviation regulations in particular. The viewing field (VF) is usually represented for clarity in the form of chart , plotted in rectangular coordinate system, wh ere and  are respectively vertical and horizontal VF angles. Openings outlines presented in rectangular coordinate system should closely comply with normative chart.

Conventional view assessment methods, including measuring ones (in-situ and virtual with 3D-model), are considerably labor consuming, as they require a lot of measured results manual processing. Besides, with the input data alteration (such as basic directional point C position) all measurements must be repeated.

The previous effort of this problem solution automation by means of 3D CAD Solid Works system resulted in successful graphic solution of VF plotting in polar coordinates. However, geometry conversion to rectangular coordinate system with the CAD system failed. To that end, an algorithm employing specially designed software was implemented. It led to the desired result. Nevertheless, the complexity of such solution makes this technique implementation quite a problem.

The objective of the effort presented in the paper was the outside world view VF plotting graphic technique with only basic 3D CAD system available in the Company.

The most challenging task while this technique developing was geometry convertion from polar to rectangular coordinates. This problem was solved in CAD NX 8.5 by successive projections on auxiliary vertical and horizontal cylinders with subsequent unfolding of these projections on plane. As a result, we have obtained a field of vision plot in rectangular coordinates. All constructions are fully associative. With input data alteration, the geometry reshapes automatically.

We managed to solve the problem of normative field of vision boundaries on the crew cabins’ surface determination, based on reverse projection combinations VF normative field of vision boundaries convolutions in rectangular coordinate system.

The main goal of the paper is determination the dependency of maximum thinning of the material from geometrical parameters during pneumo-thermal forming (superplastic forming — SPF) of various standard forms.

This paper describes the data, obtained by studying the cylindrical shape components formation.

To solve this problem, mathematical simulation of cylinders of various configurations of alloy VT20 was made with PAM STAMP 2G software package.

The simulation was performed for the groups of cylinders with radius of R = 50 mm and thickness of S = 1 mm, and with radius of R =80 mm and thickness of S = 2 mm. The results of the simulation are presented by S’R’ diagram for various H’ values.

The lines were constructed using linear approximation, and coefficients “a” and “b” were obtained for approximate lines. Also graphs a(H’) and b(H’) were constructed.

Substituting the obtained values of coefficients a(H’) and b(H’) in the of the line equation, approximate expression was obtained for estimating minimum wall thickness for AMg6M alloy cylinders with R = 50 mm, R = 80 mm and cylinders with R = 80 mm from.

As it is seen from the dependencies, these coefficients are approximately equal, which allows us to extend this dependency to other group of cylinders.

The dependencies were tested with AMg6M alloy.

Component thickness was measured by ICH-10 indicator. The experimental and simulation results revealed good convergence.

After the cylinder simulation results analysis, recommending cylinder forming selection graph, depending on the required thinning [S/S0] Ч 100% was developed. This graph presents the recommendation for designers and technologists. Its main purpose is to simplify the design of components with optimal mass characteristics. The knowledge to what relative parameters area a component belongs would sate the design time.

Thus, the following ratios for cylindrical components are considered optimal:

1. For zones with thinning greater than 78%, better apply molding or drawing.

2. For zones with thinning of 45-78% (0,1 < H’ < 0,75 и 0,05 < R’< 0,7) the best optimal solution SPF technology implementation.

3. For zones with thinning of 25-45% the better solution is to use pre-stretch for set of material from flange zone with following pneumatic forming. For zones with thinning of 16-25% (for components with H’ > 0,8 and R’ < 0,5) SPF technology implementation is not advisable.

The realization of so-called “Fully-Electrified-Aircraft” concept (i. e. without any on-board air-powered and fluid drives) assumes a significant power increase of on-board main electric generator (up to 500 kW per a single power installation with two build-in high-speed generators).

Up to date, cascaded-synchronous generators with brushless electromagnetic rotor excitation from synchronous cross-field exciter through uncontrolled rotating rectifier-exciter were traditionally used as the main electric generators (including starter-generators) with capacity range from 30 kW to 250 kW. Their essential drawbacks are as follows:

– structural complexity of salient-pole rotor design;

The cascade generator without constant rotary speed driving gear replacement by magneto-electric generator (with rotary permanent magnets) leads to the following complications:

– The necessity of implementing fast powerful emergency releaser on the drive shaft to provide armature-coil short circuit protection;

– The necessity to install a full-size (with respect to power) armature circuit static voltage regulator.

The following positive moments should be taken into account for such alternative justification:

1. undisputed advantages of a classic asynchronous motor with «squirrel cage» rotor circuit are as follows: reliability, workability, high rotation speed, small air gap, wide variety of cooling facilities, thermal stability, fair weight and size parameters and starting performance;

2. higher voltage combined AC-DC power supply system of displays decent electric energy compatibility with the voltage inverted supply circuit with higher DC voltage of 0 + 270 V and grounded mid-potential conductor;

3. synchronous compensator with permanent magnets can be used for self-excitation to unload the sine wave voltage inverter in AM-generating mode;

4. asynchronous generators with excitation from sine-wave voltage inverters together with synchronous equalizer, allow provide trouble-free parallel operation of two and even more AC power-supply channels.

The last of the above-listed factors has significant limitations. That is why combined AC-DC power supply systems without constant-rotary-speed-driving gear assume usually parallel operation only for the channels with DC (rectified) voltage (270 V or 540 V) irrespective of the types of main generators.

With such local sub-systems integration into common centralized power-supply system (270...540 V) with high specified total power (up to 1 kW) provides high quality electric power both in static and dynamic (transitional) modes.

It seems a reasonable try to realize a trade-off (at least at the present time) combined system for AC-DC higher voltage power-supply system with separate main channels of higher power generating and distribution subsystems of three types fixed in the Russian State Standard “P 54073-2010”:

1) three-phase unstable (“floating) frequency AC voltage: 115/200V or 230/400V, 360-800Hz; 2) higher DC voltage: ± 270V or ± 540V; 3) three-phase stable frequency AC voltage: 115/200V or 230/400V, 400 Hz.

As to separate classic backup low-voltage ( ± 27 V) DC power-supply sub-system with rechargeable battery regulated by Russian State Standard for low-voltage power consumers as well as backup electric-motor and/or convertor-invertor transducers — its presence, surely, is provided with any power-supply system version.

While designing up-to-date high-performance on-board electronic and electrical system distinguished by their weight-dimension, reliability, value and operation- economical parameters the preference should be given to a unified modular power-scalable architecture. At that, the best effectiveness is obtained under corresponding circuit design selection for equipment modules with reversible (bidirectional) conversion, such as, buck/boost reversible switched mode converters (direct or transformer versions – 270/27 V); reversible rectifying-inverter converters with power factor correction (115/200 V – 0 ± 270V), reversible frequency converter (360...800 Hz/400Hz) and the others as well as multifunctional pulse converters. These preferences provide rather flexible adaptivity in various units, devices, assemblages and sub-systems upgrading.

AC-DC combined power-supply system structures based on reversible electric-machine (starter-generator and engine-compensatory) and static (rectifying- inverter, inverter-rectifying and various-frequency) converters examined in this article successfully meet basic criterion that is submitted to a promising Fully- Electrified-Aircraft electrical equipment. The suggested circuit design for the main power-supply channels and converters seems to be suitable for A/M scalable systems of versatile modular architecture with high mass-energy, reliability and operation-economical effectiveness. It should be noted that such a circuit design is protected by Russian Federation priority.

Implementation of vector pulse width modulation (VPWM) for of aircraft electric motor drives systems converters becomes one of topical trends. VPWM nowadays found application in many branches of industry and manufacturing to provide the quality of various electric motor drive systems control. Depending on motor connection to the inverter method several options of developing the power stage of the inverter are possible. The paper presents the structures most frequently used in practice. VPWM realization with three-phase bridge is most expedient. To control voltages across motor windings with microprocessor unit it is necessary to connect control action with switching sequence of power switches by analytic expression. It is necessary herewith to bear in mind that power semiconductor devices switching does not happen instantaneously. To obtain optimal analytic expression we used the technique based on Karnaugh-Veitch maps. The paper presents Karnaugh maps and command word (CW) generation for a three-phase bridge corresponding to the specified switching algorithm. It shows Karnaugh map permitted transitions. To obtain necessary command word allowing through currents elimination we suggest implementation of additional vectors. The states of power switches corresponding to these vectors are also presented. The paper presets such vector generating sequences, where the through currents do not occur, as well as the states preventing motor discontinuous currents. Analysis of suggested algorithms was carried out using MATLAB. Computer model of VPWM is presented. Breadboard testing fully confirmed the efficiency of suggested algorithms. The obtained analytical expressions for switching algorithms command word generation allowed eliminate through currents, reduce speed fluctuation amplitudes and induction motor torque ripples, as well as motor drive response speed.

The development of aviation technology gradually leads to increase of the on-board electrics and electronics number, including consumers of the first category, which are indispensable for the safe completion of the flight. In case of providing power from batteries, the issue of high-grade power delivery for vital consumers arises.

As a candidate solution to the problem of inverter output power increasing we suggest to use multi-cell inverter comprising of several unit inverters, allowing distribute the total load. Besides, microcontroller implementation can result in better shape of the output voltage, increase of efficiency and weight and size characteristics reduction.

The paper considers the technique of inverter output voltage generating according to an algorithm allowing approximate sinewave signal and provide equal loading of each unit inverter. Such an approximation allows increase the output voltage quality, reduce the size of each output filter, while the uniform load will allow unify the unit inverter and enhance the system flexibility in case of one of the components failure.

For an odd number of inverter cells the amplitude of the approximated signal is identical to the reference sinewave and circumscribes the sinewave over its highest point.

To provide a uniform load of all inverter cells, the author suggests divide each time interval of approximating voltage level into equal subintervals according to the number of inverter cells. With this, each voltage level is formed as a serial and parallel addition of the inverter cells voltages.

This method of the inverter cells power switches commutation allows create scalable switched mode structures of multi cell inverters. Theoretically, the maximum number of inverter cells depends on the inverter cell power switch minimum switching time.

The paper compares the simulation results obtained for inverters with even and odd number of power cells. The variant with odd number of cells dempstrated the best characteristics despite the lower number of unit cells.

The paper presents the study of the effect on AK12 alloy structure and mechanical properties while subjecting its melt to complex modifying treatment with carbon-bearing Freon12 gas and salt compounds.

Based on earlier obtained results, we sel ected K₂TiF₆, BaCO₃ salt compounds and carbon (graphite) as modifying compounds. KA12 alloy processing with the sel ected compounds was performed both without and with its melt blowing-down by the above mentioned gas.

Mechanical tests results and microstructure analysis revealed that the Freon-12 blowing-down makes modifying impact on AK12 alloy. The highest mechanical properties, and refining of ?-solid solution dendrites and refining of silicon particles in an eutectic is observed at simultaneous treatment of an AK12 alloy by the modifying salts (BaCO₃ + K₂TiF₆ + C) together with Freon. The strength of the alloy equals to176 MPa, and relative elongation equals to 9,40%.

Besides, the Freon blowing-down increases the modifying effect for the majority of considered flux compounds. The alloy AK12 treatment with Freon and salts (BaCO₃ + K₂TiF₆) is most perspective, compared to the alloy modified only by salts (BaCO₃ + K₂TiF₆). The strength increased fr om 162 to 188 MPa, and elongation - from 7,70 to 9,01%.

Based on thermograms obtained while the melt treatment, surface morphology data and elementary composition of AK12 alloy treated with test compounds, we made a conclusion on the presence of carbon-bearing disperse particles similar in composition to aluminum and titanium carbides in threated with salt compounds (BaCO₃ + K₂TiF₆) and Freon alloy. Those formed carbides can be additional crystallization centers in the melt due to dimensional and structural resemblance with aluminum crystal lattice.

Analysis of the results of the carried on research revealed the perspective of carbon-bearing materials, and Freon 12 in particular, implementation as aluminum-silicon alloy modifiers.

Film structure of aluminum-silicon alloy (99% Al and 1% Si) obtained by physical magnetron sputtering was studied by electron microscopy and profilometry. The values of average aluminum-silicon alloy films with varying sputtering power, sputtering time and metal coating thickness were studied and obtained.

With equal aluminum-silicon alloy films thickness, the sputtering time decrease results in alloy grains average size and quantity reduction. With power increase and sputtering time approximately twofold decrease the average grain size and number of grains per square millimeter decrease proportionally. The obtained grains size depends weakly on sputtering power and stays within the range of 290-330 nm with ± 7% precision.

The average film grains size of aluminum-silicon alloy increases non-linearly approximately 25 times (from 20 nm to 500 nm) with film thickness of aluminum-silicon alloy increase from 0.2 mcm to 1.2 mcm due to sputtering time increase, which probably can be explained by substrate temperature rising due to of aluminum-silicon alloy condensation. The substrate temperature rise results in grain size increase.

The grain size and their quantity are practically independent from sputtering power, but they depend on time of continuous residing in plasma burning zone and plates cooling efficiency. In vacuum, the plates surface has no time for cooling in the process of deposition, which leads to excessive ions mobility of sputtered material, grains formation and growth. The more time the plates reside in plasma burning zone, the higher the intensity of grains growth.

Aerospace industry suppliers quality management systems (QMS) certification procedures differ substantially from each other at national and international levels. Depending on what objectives the organization requesting the audit pursue, it should decide on the type of the standard, against which to implement the QMS. It may be QMS the general standard ISO 9001, the industry standard AS 9100 or the military standard GOST RV 0015-002). It should decide also on passing conformity with QMS standards requirements certification, engaging either national or international certification body. The goal of this paper consists in considering the certification specifics at national and international levels and reveal their main advantages and disadvantages by comparative analysis, as well as work out recommendations on upgrading the certification effectiveness level at the national level.

The results of analysis revealed that the certification procedures for compliance with the AS 9100 international standard are the most strictly regulated and controlled by the accreditation bodies, regulatory bodies and aerospace industry suppliers themselves. Such a level of control of certification procedures is possible due to a number of distinctive features of audits conducting for compliance with AS 9100industry standard. Firstly, while certification conducting in the aerospace industry the ICOP (Industry Controlled Other Party) scheme functions. ICOP is a conception, which assigns еthe right to control the suppliers to the representatives of the aerospace industry, certification bodies and accreditation bodies. Secondly, the results of audits, as well as information about all the participants of certification procedures (accreditation bodies, certification bodies, auditors and certified suppliers) is displayed in the on-line database OASIS (Online Aerospace Supplier Information System), which ensures the transparency of certification procedures. In addition, the aerospace industry is characterized by the high degree of involvement of all concerned parties in formulating the requirements for conducting audits.

The listed specifics of aerospace industry supplies certification at the international level guarantee a high level of confidence in the issued international AS 9100 certificates. Accordingly, for the goals of company's development at the international level and for obtaining more customers, it makes sense for the aerospace industry suppliers to undergo certification for compliance with AS 9100 at the international certification bodies. Those conclusions could be used by Russians aerospace suppliers while selecting the certification body.

The second application area of the paper results is the Russian certification and accreditation system. As for Russian certification bodies, to improve the efficiency of their operations, as well as increase the level of confidence in the issued national certificates, an important step consists in study and use of the experience of the international system of QMS certification in the aerospace industry, as well as the Russian accreditation body (Rosaccreditation) joining the joining the International accreditation forum and the ICOP scheme.

The article is devoted to Russian corporations' competitive recovery realization in modern economic conditions. The authors study the issues of national hi-tech enterprises competitiveness by, among other things, these enterprises strategic plan indicators improvement.

Complex analysis of modern administrative methods allowing optimize the production program of the enterprises based on improvement of the researches organizational technological level (OTL), design and production was carried out.

The article defines organizational and methodological principles of optimal organization and production management system, and presents the developed model of optimal production and planning system model based on organizing and methodological principles and methods aimed at successful realization of goals and tasks of the enterprise.

The authors formulated one of such methods, namely, a method of indicators optimization in conjunction with developing production OTL as follows: the conventionsl production OTL is determined by progressive norms and regulations, and optimal OTL is suggested for a specified program. The obtained optimal OTL indicators form the basis of developing and realization of production OTL modernization from the existing level to optimal.

The authors present the method of strategic plan indicators optimization using adaptive economic and mathematical models with more details. They stressed economic efficiency of the introduced innovations while determining the enterprises' optimum production program including the suggested special standard coefficients, which adapt an abstract EMM to specific conditions of corporation development. It provides facilities for system and integrated approach while developing enterprise management strategy.

The authors have developed a scientific and production enterprise innovative projects risks assessment method.

Difficulties of identifying risks of various types and a problem of their assessment are associated with the fact that Scientific Production Enterprises realize simultaneously a set of projects being at different stages of life cycle.

In this regard, the authors offer the use the “risk portfolio”, which means the overall risk of innovative projects portfolio, which size does not exceed the appetite risk level accepted by the company.

By “risk portfolio” assessment, the authors mean the process of identification and quantitative assessment of innovation projects portfolio realized simultaneously by the enterprise.

The quantitative assessment procedure of risks stipulating the risk portfolio level is based on determining the value composing it by questionnaire survey of experts.

The electronic questionnaire in MS Excel allowing obtain an integrated assessment of a risk event after entering certain data introduced by experts has been developed to assess the risks of innovative projects realized by the scientific and production enterprise.

Thus, after risks levels assessment of the innovative projects realized simultaneously by the enterprise the possibility to determine the size of the risk portfolio becomes possible.

Modeling of innovative projects realized by Scientific Production Enterprise risk portfolio state can be realized by creating a cognitive map.

The method offered by the authors will allow control the size risk portfolio within the risk appetite level accepted by the enterprise.

The paper presents the analysis of the existing approaches to industrial parks forming. The analysis revealed that much attention is payed attention to this problem. The issues concerning industrial parks terminology and classification, their establishing criteria, structural elements, business legal structure and sources of financing are also tackled. However, mechanisms of an industrial park establishing and functioning are not considered. The industrial parks based on industrial enterprises forming assessment system, unity on terminology and industrial parks classification as well as a single mechanism of their forming, functioning and regulation on a statutable level are also missed.

It should be noted that, in some cases, industrial parks present an effective tool for solving economic and social problems, as well as contributing to the development of Russias innovative potential.

Theoretical and methodological basics of industrial parks based industrial enterprises forming are developed.

The paper substantiates the possibility and actuality of industrial parks creation based on industrial enterprises. On the one hand, the need to solve socio-economic problems of regions, on the other hand, the presence of starting conditions for the industrial parks creation (production areas, equipment, financial capital).

The paper suggests criteria allowing identify industrial enterprises as the territories with potential for industrial parks establishing potential.

The purpose of industrial parks creation based on industrial enterprises consists in improving the competitiveness of products through the introduction of science-intensive technologies; diversification of production; development of knowledge-based industries; development of innovative entrepreneurship and infrastructure. The paper suggests the industrial park organizational and functional structure, including research and development block (research establishment, laboratories, and experimental design bureau), production unit (enterprise, manufacturing a product), infrastructure block (firms, engaged in financial, marketing, legal and other services).

Industrial park fiscal system is suggested.

The paper reveals the most important criteria for industrial park functioning:

• Location.

• Property complex size.

• Total land plot area size.

• The size of enterprise production area, including uncommitted.

• The presence of innovation potential.

• The presence of high-grade in their spheres specialists.

• The presence of high-grade conditions for possible resident companies allocation.

Modern advanced helicopter should have not only vertical take off and landing and hover capabilities but also high cruise speed and long range. Compound helicopter can reach high speed with the same engine power by optimal use of the main rotor, additional fixed wing and propulsion system. Producing lift wing reduces main rotor thrust required and prevents the retreating blade stall. Propulsion system thrust provides the fuselage optimum angle of attack when it does not produce negative lift and has minimum drag.

One of the most effective ways to produce thrust at high speeds is tilting the tail rotor. In this case all the power from the main turboshaft engines will be used at hover and also at high speeds, shared between main and tail rotors.

As an example the authors take single rotor helicopter with two VK-2500 turboshaft engines, additional 15 m² wing, propulsion prop with 3.2 m diameter and 11900 kg take off weight. To reach maximum cruise speed (385 km/h) 2500 hp should go to the tail propeller which is 70% of all power produced.

However, different blade washout is required at cruise speed flight (the longest flight mode) and hover for best performance of such a propeller. For this purpose at Mil Moscow Helicopter Plant flight simulator the most effective washout was determined.

Because of additional wing and tail rotor tilting system compound helicopter is 400 kg heavier then a conversional one. But even with lower service ceiling (greater weight, negative wing lift at hover and more anti-torque rotor power required) it has 15% higher cruise speed and range which can be shown both at high and low flight altitudes with up to 13000 kg takeoff weight.

The article considers a method of statistical synthesis of a distribution function of targets between various types of aircraft. It presents statistical samples, describing an optimal type aircraft system in large, as well as samples describing the function of target distribution.

The process of building of design and functional bindings that model aircraft systems at large starts from forming a statistical sample, which input data represents the characteristics of a goal set, while elementary functions of targets distribution by aircraft types are taken as intermediate characteristics. We accept the values of criterion of optimality as output data. This work assumes as criterion of optimality the cost of aircraft system. Intermediate characteristics have special meaningfulness when modeling aircraft systems.

According to the principle of mathematical model self-organization, at successive complexity increase of a model (in the course of the transition from a linear model to a square one and further to higher degree models) all external criteria pass through their minimum. It gives the possibility to obtain a model of optimal complexity, unique for each criterion. In self-organizing theory of complex systems models all basic algorithms for building models of optimal complexity are based on grouped account of arguments. Combinatorial algorithms appear effective for problem order no more than the specified sample size. For problems of higher complexity, such as designing systems for an optimal type aircraft, the multi-row algorithms become more effective. The perspective of multi-row algorithms should be noted, since, in principle, they are the prototypes of genetic algorithms, which allow solving simulation problems of with dimension of several thousand variables. However, whichever the algorithm is, it does not allow going beyond the framework of the specified class of basic functions. These algorithms increase only the complexity of the model within a specified basis.

With statistical synthesis employed in the paper the modeling algorithms are built in the way that provides the possibility in principle to obtain the models involving various basic functions.

This corresponds to the basic statistical synthesis concept, according to which the output data of the initial statistical sample is modified in a certain manner in the process of building of a mathematical model. This adjustment is carried out according to the conditions and requirements that the formed model should meet. Thus, at each stage of the mathematical model building we form the statistical sample inherent to it, and the inherent optimal system of basic functions corresponds to each sample.

The article presents the analytical models of the targets distribution function, which represent an approximation of statistics reported in terms of trigonometric polynomials.

The article considers the operation of statistical sampling reduction, which reduces the initial n — dimensional sample to n one-dimensional samples, and operation of inversion allowing obtaining the inverse sampling required for the formation of inverse functions. Based on these operations, build one-dimensional functional relationships between characteristic functions and outer target set characteristics presented in the form of trigonometric polynomials. The paper presents a simple, but at the same time effective way to meet the statistical functional limitations. This method is based on statistical sampling regularizing. It consists in replacing or eliminating fragments of statistical sampling, which do not agree with the specified limitations.

Aircraft dynamics modeling at high angles of attack requires adequate reproduction of all forces and moments acting on the aircraft under these conditions. Until recently, accounting for the ambiguity (non — stationarity) of lifting force and pitch moment coefficients c_y and m_z dependencies causes difficulties due to the lack of simple enough mathematical models of this phenomenon. The problem is aggravated by the need to calculate the aircraft dynamics in real time, which is mandatory for the bench research and crew training simulators.

The paper considers the problem of developing mathematical model of such kind based on the data obtained by the results of the tests of the of twin-engine passenger plane of the traditional scheme model with the of OVP-102B installation, which reproduces the aircraft model angle of attack harmonic oscillations at different setting angles and frequencies.

Based on the analysis of dynamic dependencies c_y(α,ά) and m_z(α,ά) obtained as test results we proposed a rather simple model for the coefficients c_y and m_z calculation as a function of the current angle of attack α(t).

The model is based on the static dependencies of lifting force and pitch moment coefficients c_y and m_z conversion by the first order differential equation (aperiodic link) with time constant T depending on the angle of attack and its rate of change.

The paper presents the technique of the input parameters and time constant forming based on known c_y(α) , m_z(α) and α(t) dependencies.

The comparison of non-stationary characteristics computed by the model with the results of wind tunnel and flight tests confirmed the suitability of the developed model for computation and experimental research of aircraft dynamics at high angles of attack.

The purpose of this work was to study the primary probe holder of a cylindrical Langmuir probe relative size effect on the results of local plasma diagnostics.

The primary probe holder radius, that should be far less than the electrons free path average length was the main subject of consideration in the presented paper, because it is the primary probe holder that is able to decrease the concentration of electrons and change other parameters of radiated plasma in the spots of its probing. It significantly affects the quality of local plasma diagnostics.

Three types of cylindrical probes made of tungsten thread of 0.15 mm in diameter were used. All three probes were provided with probe holders of the same diameter of 1.6 mm. The idea of the work consisted in obtaining measurements from the probes of various lengths under equal impact of the primary probe holder, which increased surface, allocated near the probed area contributes to recombination of charged plasma particles on its surface, and, thus reduces the level of ionization equilibrium in the radiated plasma. The cylindrical probe of a certain length averages the parameters of plasma in its scope. The invariable local distortion of plasma parameters near the probe holder affects differently the measured results for probes of various lengths. The work employs such probe lengths range that allows point out their bound, outside which the disturbances of the considered type become less than the total error of probe measurements.

The paper describes the experiment with HF induction (HFI) discharge in Xenon and probe measurements carried out using probe station Plasma Sensors VGPS-12. A number of technical features of VGPS-12 allows increase the accuracy of plasma diagnostics results, narrowing down, for example, electron density measuring error field to the value of about 10%. These features include: implementation of Dryuvesteyn method that does not require prior guesses on the shape of the electron energy distribution function in contrast to the other probe techniques; suppression of the most part of errors by protecting reference electrode with developed surface; probe surface cleaning by ion bombardment and HF current heating.

The work is concluded by the analysis of the obtained results, allowing formulate recommendations on selection of the main sizes cylindrical Langmuir probes to ensure the acceptable plasma diagnostics accuracy.

The article presents the results of experimental determination of the of the piston engine share in the total power plants noise of the light propeller aircrafts An-2 and Yak-18T, MAI-223M and MAI-890U, performed under static conditions at the local aerodrome. The article provides a brief overview of the mechanisms of noise generation by the aircraft piston engines. The power plant emission band consists of harmonic and broadband components. Its sources are piston engines and propellers.

Based on measurements at several points of the acoustic far-field narrowband spectra we separated harmonic components emitted by the propeller from those emitted by the engine. Separation of the high-frequency component of the engine broadband noise against the background of the propeller whirligig noise appeared to beimpossible. The possible source of the dominant radiation is the turbulent wake behind the rotating blades. The important share of piston engine harmonic emission of the total emission power of engine-propeller power plant is experimentally revealed, and under lower engine operating modes An-2 and Yak-18T (at Mach circumferential propeller velocities less than 0.7) in particular.

The paper presents the factors affecting the piston engine share in the total propeller power plant noise. It also studied the effect of nosing engine on acoustic power level of the engine fundamental tone. It was found that the nosing of the engine could be considered as one of the waysof the afield noise reduction of ultralight aircraft MAI-890U.

We obtained acoustic emission patternsof piston engines typical for light aircraft. Aircraft ASh-62IR and ROTAX-912ULSengines emissioncharacteristic maximums correspond to the azimuth angles of 0° in the forward hemisphere and of 135-150° in the rear hemisphere. M-14P engine emissionis relatively even over the space in the direction angles of 60-120°. Typical minimum levels of engine noise observed on the axis of the crankshaft, i.e. in the direction of 0° in the forward hemisphere and 180° in the rear hemisphere. One can use these emission patterns in the future for the noise prediction models of aircraft piston engines.

The paper tackles the issue of effective space system design. The experience of developed countries proves that such transportation system may include hypersonic loop. The Faoulty of Flying Vehicles Engines in Moscow Aviation Institute (National Research University) has developed the conception of the reusable space transportation system (STS) based on liquid rocket engine (LRE) and scramjet.

The idea of using available stock of propulsion systems development forms the basis of the concept. Space Transportation System consists of two stages: the booster (first) stage and the orbital (second) stage. We plan to include RD-161P engine into the booster stage of the propulsion system. This engine is in progress at «Scientific-production association „Energomash named after academician V.P. Glushko“».

Orbiter has a dual-loop propulsion installation consisting of a scramjet and a low-thrust thrust (500 N) rocket engine. These engines are developed at the Faoulty of Flying Vehicles Engines in Moscow Aviation Institute.

To select the fuel type for the use in space transportation system, we compared several fuel compositions that meet the environmental requirements, high efficiency and assimilation. As a result, we chose the fuel: high-concentrated hydrogen peroxide and kerosene.

The successful flight of a returnable STS requires taking into account the effect of flight conditions on the flying vehicle control. We propose to create a demonstrator and simulate all phases of flight STS with the aircraft of significantly smaller size.

The following demonstrator flight scheme is assumed: after orbital stage separation, the first stage continues moving along its ballistic trajectory. Parachute is ejected, and then at the height of 3km helicopter grabs the first stage. The helicopter delivers dry assembly to the launch pad for reuse.

We conducted ballistic calculations and preliminary design studies. As a result, we obtained weight characteristics and the model of a demonstrator. Payload weight is about 4% of the take-off weight of a

demonstrator. With an increase of take-off weight of an aircraft up to actual STS weight levels (100-200 tons), the aircraft placing into LEO payload efficiency reaches 10% of the take-off weight.

Using well-known and calculated demonstrator performance we developed demonstrator flight control system: ground control post hardware and on-board equipment.

Phases and modes of flight demonstrator are similar to space station Buran project. It facilitates the implementation of the existing stock and flight control algorithms.

The developed concept of a two-stage reusable space transportation system is designed with two types of propulsion — LRE and scramjet, operating with high-concentrated hydrogen peroxide and kerosene. We suggest to use the demonstrator to make more precise the impact of flight conditions on the efficiency of STS. Calculations кreveal that the efficiency of the aircraft payload placing into LEO reaches 10% of the take-off weight.

The subject of the research is ultralow power turbine with axial nozzle set and diagonal impeller. The purpose of this work consists in improving the ultralow power turbine efficiency.

The paper describes the features of the work process in ultralow power turbines with diagonal impeller based on the equations of momentum and energy. It also studies the effect of the impeller degree of bias and the ways of its meridian profiling on energy characteristics of nozzle set, the impeller and the turbine in whole, using numerical methods for viscous flow 3D modeling with ANSYS CFX. The results of the analysis are based on theoretical concepts of work process within the diagonal turbine impeller.

The author derived equations of flow radial balance and flow energy within the impeller of diagonal type. It was found that implementation of diagonal impeller rather than axial provides an effective solution. It allows increase of power efficiency of the turbine. The turbine efficiency at rating conditions with loading implementation of diagonal impeller rather than axial parameter Y_T = 0,23, the rate of pressure reduction and calculated degree of reaction was increased by 9% due to diagonal impeller implementation with constant blade height in meridional section with impeller midline to turbine axis tilting angle . It was found also that impeller meridional profiling technique does not affect turbine effectiveness within the range of rational values .

Small size turbines with shaft power N_t = = 0.01...10 kW and working fluid flow are meant for various accessory drives and auxiliary power plants, energy systems for space and surface transport, as well as processing tools.

1. Implementation of diagonal impeller rather than axial allows power efficiency increase of about 9%. The highest efficiency increase is achieved with impeller midline to turbine axis tilting angle values . 

2. The method of the diagonal impeller meridian profiling does not affect the turbine efficiency within the range of , so it is recommended to use the most workable method of profiling with h_{i imp} = const. The results of the study display a significant effect of diagonality on impeller work process, which stipulates the necessity of further theoretical and experimental studies of the diagonal turbines and turbines with diagonal impellers of various types.

The goal of this work consists in substantiation of creating the central maintenance and repair (M&R) system at Myanmar airports, and development the criterion of its breakeven.

The author carried out retrospective analysis of aerodrome network and air transport in Myanmar. The conclusion is drawn about the necessity of creating technical maintenance and repair system, one of the main elements of integrated logistics support. The criterion, which must be followed while forming of such kind of system, was developed, and this criterion is system operation breakeven. This means that the M&R system total operating revenue over a calculation period has to be greater than M&R system total expenditure over the same period.

Since the volume of proceeds is determined by the market of such type of services, minimization of M&R system costs should be the basis of this problem solution. This solution requires determine the compromise between aircraft designers (consideration of M&R system personality for each aircraft) and interests of air companies, reaching after reducing aircraft M&R costs upon observance of flight safety requirements. Creation of such network requires scientific substantiation given in this paper. The obtained results can be used while taking investment decisions related to the development of Myanmar economy.

The deficiency of aircraft M&R may entail drastic consequences, such as poor condition of aircrafts and, consequently, increase in aviation accidents. It will lead not only to deterioration of air companies position domestically, but also will seriously affect the budget, for tourism is the most important revenue item of the country. The paper considers the problem of aircraft maintenance and repair at Myanmar airports. The paper reveals the deficiency of such system in Myanmar. For such system formation, we need to elaborate requirements to M&R system and Myanmar airports with allowance for airlines interests, form general criterion and constraint system. On their basis we should develop the maintenance and repair system for Myanmar airports.

This article presens the analysis of the of expert evaluation methods (EEM) implementation for planning ground development test (GDT) of rocket technology products. To do this, all necessary conditions for experts’ qualitative evaluation were considered and recommended EEM were sel ected. Multilevel EEM implementation is marked herewith, where the upper layer determines the order of experts’ polling — discussion (brainstorming, synectics, commission of experts and court), or questionnaire (Delphi method). While selection of low-level EEM is determined by the specific practical problem. To simplify GDT optimal planning all products are sorted according to qualification status (corresponds to qualification method). Then selection of types, test categories to develop the required operation characteristics and their sequence for each assembly unit is carried out (corresponds to scenario methods or PATTERN, forecast graph). The problem of selection and assortment of the necessary types, categories of tests is just the most complex one during GDT plan elaboration. It is suggested to solve this problem by the selection of the appropriate product-analog (corresponds to Churchman-Akof method) with allowance for its performances, as well as a product (assembly unit) requiring optimization. This will reduce the duration and cost of GDT planning phase.

Further development of the assembly unit optimization plan for those operation characteristics that differ fr om the characteristics of the prototype, types and test categories are determined by any of the recommended EEM (mostly by scenario methods and brainstorming). Thus, as a result of the multi-level application of EEM, it is possible to elaborate an experimental development optimal plan in terms of minimization of material and time resources.

The article considers the database (DB) designed for software complex of automated flight planning for unmanned spacecrafts (SC) [1]. This DB is meant for storing data for flight planning complex, as well as exchanging data with other complexes of the Mission Control Center (MCC) and the external organizations involved in the spacecraft control process.

The method under consideration for developing a database involves the use of a relational database architecture. A relational database presents all data in rows and columns of the table. Areas of tables allocation are called table space. Apart fr om these tables they also contain indexes, notations, constraints, rules, default settings, triggers, stored procedures, user-defined functions and data types. The above listed elements are the objects of the database that are used for its building. The unmanned spacecraft automated flight planning database is structurally divided into four table spaces. The first contains information about the space-time state of the spacecraft. The second contains data necessary to solve the problems of automated planning and its automated checking procedure. The third is the archive of data exchange between software complex of automated flight planning and MCC complexes. The fourth stores output data, transport files and receipt tickets generated in the course of the exchange with the MCC complexes and external organizations involved in spacecraft control process.

All DB tables are arranged in accordance with hierarchical structure. We use two types of tables: table header and their subordinates. The header table contains general information and characteristics of the object, such as the composition of spacecraft groups or ground stations that provide control, while subordinate ones contain specific data structure of the spacecraft equipment, lists of commands, the types of tools used during the sessions.

The database has been designed to integrate the individual elements of unmanned spacecraft automated flight planning into a unified complex. It presents the link between the complex software modules and provides their interaction with each other and with MCC complexes.

The outcome of the work led to the development of unmanned spacecraft automated flight planning software complex Database. The proposed structure of the database allowed arrange data, set forth its hierarchy, provide storage and access to data and support data integrity.

The cre ate d database allowed simplify the separate software modules interrelation procedure of the automated flight planning complex and MCC complexes, providing multistream data access as well as simplifying the search for necessary information.

The developed Database was implemented at MCC as a part of an automated spacecraft flight planning software complex. It saw used in the course of flight control of a spacecraft of scientific and socio-economic purpose.

The automated flight planning for unmanned spacecraft software complex Database can be used for automatic spacecraft of scientific and socio-economic purpose control.

1. The developed database is the binding element of an unmanned spacecraft automated flight planning. It helps to access and store the software complex data and to maintain its integrity as well.

2. This relational database architecture has allowed develop the structure wh ere all automated flight planning data was divided into table spaces. Their objects are tables formed by levels and hierarchy, notations, stored procedures, triggers and indices simplifying the searching procedure and data entry.

3. Using the database also made it possible to simplify the procedure of interaction of automated flight planning software modules, providing them with data access independently of one another. This significantly simplified the design of the software complex itself, allowing its construction with separate independent blocks.

4. The Database was implemented at the MCC and was used for spacecraft of scientific and socio-economic purpose automated flight control as a part of flight planning complex.

The capabilities of modern photovoltaic transducers form the performances of the solar energy powered UAV. Solar energy ensures the flight at low speeds. High-altitude flight allows avoid the impact of clouds and provides a large view scope. High-altitude flight is carried out within the high kinematic viscosity environment that significantly reduces thereby the Reynolds number. An important feature of the aircraft consuming solar energy to maintain its flight consists in specific range of flight heights and speeds, characterized by low Reynolds number. These flight conditions are associated with non-linear dependence of aerodynamic characteristics from the incidence angle, which complicates the use of traditional methods to optimize the flight path. The flying vehicles of such kind require the technique allowing optimize the flight path with allowance for these features. The proposed climbing trajectory optimization method partitions the optimization process into two stages. The first stage consists in flight parameters region characterization, guaranteed to exceed the specified value, while the second stage is the stage of trajectory providing minimum power consumption search. The search for the trajectory allowing minimum power consumption is carried out by the direct numerical method, without characteristics linearization. Such an approach allows optimize climbing trajectory of the aircraft having non-linear aerodynamic characteristics, peculiar to the flight at low Reynolds numbers. We obtained climbing trajectories with allowance for Reynolds number and without it. The results show that energy consumption while climbing with allowance for non-linear aerodynamic characteristics are about 4% higher than the results obtained without the regard of those non-linarites. This may cause the energy shortage for flight support and lead to multi-day mission failure. Optimization of high- altitude solar energy powered UAV flight path requires regard of Reynolds number effect of aerodynamic characteristics. The flight paths computation of the UFVs of such kind is worthwhile to carry out by numerical methods, stable to aircraft aerodynamic characteristics non-linarites.

The article presents the structure of the on-board radio-electronic equipment that meets modern international requirements of CNS/ATM conception and is based on integrated modular avionics (IMA) principles. Implementation of IMA conception allows provide high reliability, enhanced functionality and compliance with modern requirements to on-board radio-electronic equipment. Architecture of the on-board equipment includes computer complex, description of information for flight deck information management field and software applications.

The computational part of the complex project includes the contents and datasheet of computing blocks, as well as the contents and designation of mezzanine modules, which perform special functions and determine functionality of the crates.

The flight deck information management field arrangement corresponds to modern allocation schemes of information consoles and allows provide their effectiveness and ergonomics. This field consists of five 15’’ multifunction displays, two head-up displays and two Electronic Flight Bag (EFB) data tablets for the pilot and co-pilot. The article presents the description of and information frames distribution over multifunction displays.

The article outlines the on-board equipment functions executed by software applications while IMA conception realization. Functional software realizes all modern communication, navigation and surveillance functions. The software applications are executed by the computation complex crates. The article suggests the applications distribution over the crates according to their functionality.

The proposed project of on-board equipment complex differs from the existing ones in large state integration of IMA systems, realizing functions of automated control systems, flight-control-navigation and radio-communication equipment. The project of on-board equipment complex for prospective transport airplane presents scalable multifunction fault-tolerant complex corresponding to the requirements for transport category aircraft, and meeting the prospective ICAO requirements to air traffic navigation and control.

Discrete weld bonding are zones of increased stress concentration, formed due to heating non-uniformity during welding process and significant difference in mechanical characteristics of a metal in weld junctions from those of a parent material. Under repeated loading action the occurrence of fatigue fractures evolved between weld materials at the contour of a weld point is most probable. The fatigue endurance of the structures with weld bonding is determined, for the most part, by the stressed state dynamic characteristics at local zones of weld points. To obtain the proper durability estimation of such kind of structures one should define stressed state dynamic characteristics with allowance for its essential spatial non-uniformity and local changes in material properties.

The paper presents the technique and results of a structure with spot-welded joints subjected to random vibration probabilistic characteristics computation using finite elements simulation. Simulation of the dynamic stress state was executed for the structure under random kinematic loading conditions with the specified power spectral density function of acceleration. To substantiate the correctness of the simulation of spot-welded joints the authors studied by computation the samples of spot-welded joints, tested for static strength under tensile and shearing, and plotted vibrating stress spectral characteristics for various points of the structure. Zones and levels of maximum stresses were determined. Features of dynamic stress state in welded joints were revealed. Estimations of specific damageability and mean longevity of welded joints under various theories of accumulation of fatigue damages were obtained.

The attribute of the study consists in detailed modelling of dynamic stress state at welded joint considering significant changes of the properties of the parent material over spot weld cross section. The properties of the material at spot weld local zones are determined based on micro-hardness yield stress empirical dependences.

The developed technique and numerical simulation results can be applied to assess the vibration strength of thin-walled structures with discrete welded connections.

During the last three decades, the design effort is being concentrated on realization of more all-electric aircraft concept. We can assign to this class the planes incorporating, either power hydraulic or pneumatic actuators for energy-intensive installations control, or electric motor drives controlling various functional objects, such as steering rudders. Nowadays the research effort is focused on two types of drives design — electromechanical and electric hydrostatic drives. The electromagnetic interference (EMI) generated herewith by electric drive motors in the form of electromagnetic fields may affect the operation of aircraft on-board equipment which, in some cases, causes failures. Furthermore, electric drive modules represent receptors of external radiated and conducted interference, which may distort control signals, formed by microcontroller and, hence, conducted EMI in motor windings in the form of phase currents transients. Thus, it is necessary to know the levels of motor phase currents, occurring during various transients, to provide normal operation of the dive and evaluate electromagnetic environment within the aircraft interior.

The goal of the paper consists in the analysis of the phase currents transients occurring in the steering drive system (SDS) motor, based on the results obtained by computer simulation in OrCad 9.2.

The paper presents waveforms of phase currents in motor windings, obtained with OrCAD simulation. The phase currents levels in actuating motor windings may be about 1.5 times greater than their rated values while operating mode variation. The above said currents and the EMI they induced in the form of electromagnetic fields affect the electromagnetic environment within the interior volume and are of serious hazard to electronic equipment of either SDS, or to aircraft on-board equipment.

The presented paper is a part of research on calculation and simulation of electromagnetic interference caused by the transients in current amplifier of an aircraft SDS motor.

The article touches upon the methods allowing increase the precision of non-rigid cylindrical parts processing.

When producing non-rigid component parts, the work pieces processed surfaces deformations subjected to the forces are scaled to machining tolerance values, which leads to the occurrence of processing errors related to it.

The article tackles the issues of finding deformations of non-rigid work pieces walls during mechanical processing, depending on cutting operating conditions, to optimize the manufacturing process and enhance accuracy.

We suggest calculate the work pieces surfaces deformation under the impact of the cutting force by using the finite-element method and the elasticity theory provisions.

As a check on the possibility of using the finite-element method to calculate deformations while processing non-rigid cylindrical work pieces, it is necessary to have an analytical checking solution. We examined such solution based in the general thin-walled shells design theory.

The paper consists of three main sections.

The first section presents a detailed calculation of deformation due to cutting force impact while non-rigid cylindrical work pieces processing according to the general thin-walled shells design theory by the example of a thin-walled cylinder. The necessary reference data to determine cutting operating conditions are included.

The second section deals with the calculations based on the finite-element method. It gives appropriate recommendations to improve thin-walled parts computational accuracy and demonstrates the graphic solution results obtained with Abaqus and Ansys programs. This section considers the selection of finite elements for thin-walled machine parts calculation as well.

The calculation results are compared with the results obtained by the exact analytical solution.

This section considers the possibilities of calculations based on contact method of cutting force estimation and other cutting parameters with Abaqus and Ansys program module. A conclusion on poor computation accuracy and impossibility of its implementation for the problem under consideration solving was made.

The paper gives recommendation, allowing correction of technological process parameters at the stage of process design, with due regard for non-rigid work pieces processing errors due to their surfaces deformation caused by cutting force. It considers the possibilities of the presented computations automation based on the finite elements method for various kinds of geometry and design features of component parts.

Experimental studies of properties of radio absorbent materials and coatings to solve the problem of aviation complexes radar visibility reduction hold a unique position. Foreign and domestic proves that without development of laboratory and testing site base achieving sound results on this issue is impossible. One of the main problems of experimental research within the framework of the problem of the radar signature reduction is the study of physical parameters of radio absorbent materials.

The purpose of the research is to develop a new measuring technique of radio absorbent and composite materials complex permittivity.

The method consists in measuring signals in orthogonal channels of a receiving device (special horn antenna) as the ratio of voltage and phase difference. For a specified voltage ratio and phase difference, we introduce such notion as “phasor”. Phasor modulus and its phase angle characterize real and imaginary parts of a material complex permittivity. This technique demonstrates improved accuracy due to account for the imaginary part of permittivity.

A special horn antenna allows increasing the penetrating power of electromagnetic waves and reduce the “effective” radius of the field interaction zone with the material. The antenna consists of a dual H-waveguide, two orthogonally arranged two receiving and two transmitting dipoles.

The paper presents measuring and computing system implementing the abovementioned method. The system consists of measuring system, mathematical support and software.

The results of experimental research of different dielectric materials, including samples of radio absorbent and composite materials are presented. Experimental studies confirm the high accuracy and sensitivity of the developed technique.

The “phasor” technique can be effectively used for studying physical parameters of existing and prospective models of radio-absorbent and composite materials for aviation complexes.

The purpose of this work is to study and compare the characteristics of electrochromic devices based on inorganic materials, such as WO₃ и Fe₄(Fe(CN)₆)₃ for using in passenger cabin viewports with dynamic shadowing.

The study of various electrochromic devices structural variants based on these materials revealed that the structure configuration based on Fe₄(Fe(CN)₆)₃ allowed obtain the most effective result. Moreover, device characteristics such as the transition time to colored state and the transmittance in the bleached state deteriorate with increasing thickness of the electrochromic layer due to deceleration Fe3+ ions reduction to Fe2+ ions. The optimal variant of electrochromic device was the device with the electrochromic coating thickness of 400 nm, which transmittance in the visible range varies from 68 to 21% when the potential difference of 1.5 V was applied for 3 seconds. With the electrolyte concentration increase, the coloring / bleaching time was found to be increased due to reduced velocity of the ions, caused by the increase of the Coulomb interaction between the ions of the electrolyte. The best coloring / bleaching time was obtained by using 1 M KCl solution.

Ease of operation, low power consumption, and high-speed dimming should be noted when such devices are implemented for passenger cabin portholes dynamic shadowing. The major shortcoming of inorganic electrochromic devices today is the limited number of coloring / bleaching cycles. Due to this, the inorganic electrochromic devices are almost completely ousted from the market by electrochromic devices based on liquid crystals and polymers (Pdlc). A disadvantage of the devices based on liquid crystals (Pdlc) is that they have only two states: fully transparent and fully shadowed, which eliminates the continuous adjustment of the device. Thus, electrochromic devices based on inorganic materials have a great potential for development and competition in the market of electrochromic devices, one of which is passenger cabin portholes with dynamic shadowing.

This work is devoted to the study of activated graphite structural characteristics and coatings on mineral fibers. It considers metallized coatings obtained by fabric vacuum spraying-up as an alternative to carbon coatings, and examines their properties, advantages and disadvantages.

By the time variation of the preliminary grinding, temperature and oxidation time the optimal conditions of chemical activation of graphite powders and particles with high particle diameter to thickness ratio were determined. Physical characteristics and thermal resistance of coatings on basalt and glass fibers were determined using various techniques of colloid graphite particles aqueous and alcohol dispersions current-conducting material coatings. Basalt fibers over glass ones substantial advantages, such as adhesion reduction due to acid nature of graphite activation, were revealed herewith.

By X-ray structure analysis and laser diffraction technique, we determined the original graphite pastes characteristics and found that graphite particles sizes and anisotropy parameters had decisive impact on the possibility of thermal shockproof radar absorbent materials formation. The X-ray structure analysis data allows make the main conclusion confirming the well-known fact that in the process of graphite powder high temperature activation the structural compression of a material and partial burnout of amorphous component occurs. According to coatings microstructure study by scanning electronic microscopy we can conclude that the films are formed by carbon particles agglomeration (flaking plates).

The paper considers various types of chemically activated graphite, which have vital structural and morphological differences. They can be formalized by the relationship of a flat particle diameter to its thickness. The possibility of obtaining thermal shockproof coating by impaction of particles with high anisotropy, i. e. planar size to thickness ratio was revealed.

The volumetric radar absorbent carbon materials of a composite fiber texture studied in the paper, which are either thermal shockproof or stable to high power electromagnetic emission, can be implemented for protection and electromagnetic compatibility provision of airspace electronic equipment. Radar absorbent carbon materials are stable under conditions of vacuum and radioactive impacts.

The purpose of the paper is to disclose the potentialities for improvements of supercapacitors, or so called ionistors, characteristics — one of the most promising types of energy storage devices along with expansion of their implementation area in aerospace engineering.

Based on the analysis of the theoretical and experimental research results published by the developers in this area, as well as original studies, the authors present several ways of improvement of supercapacitor characteristics, first of all, their charge capacity and accumulated energy. It is proposed in particular to optimize the structure and the material of the electrode. For electrodes based on nanodispersed layered graphite structures the authors show the necessity and possibility of the availability of nanoporous electrode surface for electrolyte ions consideration.

Another approach for increasing capacitance is a rational choice of an electrolyte. Different variants of aqueous and non-aqueous electrolytes, as well as solid electrolyte are analyzed in the paper. Advantages and drawbacks of various types of electrolytes are shown. To increase the energy accumulated by ionistors as well as extend the voltage range it is proposed to use non-aqueous electrolytes and create «nonsymmetrical» ionistors with redox process involved. Experimental testing of identical supercapacitor cells with different electrolyte solutions and their mixtures showed that the mixture of acetonitrile and ethylene-carbonate provided the best set of supercapacitor parameters (specific capacitance, self-discharge resistance and series resistance).

The authors carried out additional testing of supercapacitor cells of various constructions with standard domestic electrodes and electrolytes to evaluate the rate of degradation process. Supercapacitors with multilayer axial structure demonstrated the highest parameter stability.

Comparative analysis of superacapacitors characteristics produced by Russian and foreign manufacturers was carried out.

Possible areas of supercapacitor implementation in aerospace engineering were studied with allowance for the peculiarities of their characteristics (high specific power and relatively small discharge time).

Supercapacitors can be effectively used for various short-term power applications, drives, etc. in combination with other power sources. We suppose that developing power sources for unmanned aerial vehicles can be the most efficient area of implementation of superconductors.

The purpose of this work is to define the role of organizational and resource support for innovative development of defense industry complex (DIC) enterprises, and rocket engine industry in particular.

To achieve the purpose the authors analyzed the DIC contribution to innovative development of the country, identified the main areas of DIC current and strategic development, and revealed the key aspects of leading enterprises innovation development.

Currently the Russian Federation faces dual challenge, namely, launching innovative development mechanisms and raise a level of national security in military, economic, technological and other spheres. A solution of the above-mentioned problem is possible through intensification of development of defense industry complex, which concentrated considerable innovative potential. Implementation of this potential can activate innovative processes in the economy in large.

In our opinion the main trends of DIC current and strategic development, are as follows:

— firstly, to determine the most efficient trends of innovative development of high-tech industrial production and further search for the reserves and factors stimulating DIC innovative development and national economy as a whole;

— secondly, to improve resource support, i. e. the system of coordination of legal, organizational, technical, financial and economic, scientific, technological and human resources at all levels of governmental and business activities.

Resource support should be based on its consideration as a system, i. e. a group of elements (financial, labor, material, manufacturing resources etc.) organized in such a way that they are able to interact in block to achieve certain goals.

The paper also analyzed the activity and the main documents regulating innovation activities at one of the leading rocket engine industry enterprises — NPO “Energomash named by Academician V.P. Glushko”. The analysis revealed that within the framework of unit power rocket liquid propellant jet engines for primary and second stages world market sector “Energomash” products has no business rivals. In this respect, the main goal of the innovative enterprise development means improving the competitiveness and economic efficiency of “Energomash” activities, guaranteeing the status of the enterprise, as one of the sectorial leader in the world market. This goal can be achieved by proper resource supporting of Innovative Development Program. “Energomash” personnel potential development is defined as a key factor of innovative development of the enterprise. The authors herewith analyzed the current personnel structure and revealed the tendencies of personnel support of the enterprise.

Under conditions of innovative economy development scientific-and-technological advance, information technologies, wide spread occurrence and implementation of innovations become the main factors of social and economic development. Innovative orientation is an integral characteristic of modern economy. Factors affecting innovative activities are miscellaneous, but cooperation of enterprises with different companies, research organization, corporations and development institutions should be noted. Highly professional team and a new product are of great importance.

Innovative activities of an enterprise should be considered in two ways: on the one hand, it is the result, presented in the form of new products or technologies; on the other hand, it means introduction of products, principles and approaches replacing the preceding ones at the enterprise. It leads to lower production costs, considerably improves consumer properties and quality of the product, and allows meet the growing needs of buyers in the market.

The effectiveness of aviation industry enterprises activities does not consist in savings on the scope and the search for inexpensive resources, but in timely responding to the changes in the external environment. Only the enterprises that are able to adapt their internal environment to external changes without detriment to their activities can hold their competitive position in the domestic and world markets.

The article discusses the issue of maintaining competitiveness through adaptation measures to changes in external environment. The efficiency of the processes during such changes at the enterprise depends on presence of its potentialities, which are determined by production, financial, investment, marketing, personnel, information, resource and scientific and technical potentials.

We define the specific features of innovative activities implementation for enterprises of aviation industry. This article presents the example of organization of conditions for for linear friction welding innovative technologies implementation at the enterprise PJSC “Ufa engine industrial Association”. The paper considers the system of innovative development of an enterprise, as well as possible improvements and the impact of external and internal factors in the process of implementing innovative technologies.

The main conclusions consist in the following: under conditions of the external environment dynamism, developing enterprises must constantly perform internal restructuring to adapt to it, to remain competitive during the strategic period. With that, the project review at various options, and considering all risks is necessary.

Realization of the system approach to cooperation of a parent enterprise of space equipment developers to the rational solution of multidimensional applied problems of the organization and carrying out international cooperation in space activities from positions of development, production and operation of space equipment is considered.

The attention is drawn to the feature of the international space cooperation associated with the need of creation of conditions for its carrying out in all spheres of the state activity. Not only scientific, technical and economic, but also political, defense and social spheres are involved.

For support of rational decisions adoption in such a complicated and multilevel process of options formation, their productivity forecasting and selection of rational decisions, one needs to invoke scientific methods of the system analysis.

At the same time, the problem of non-systematized use of numerous system methods that negatively affect the adequacy of decisions by the target is revealed.

Based on experience of the parent scientific and production enterprise engaged if creation of complex system developments and its participation in MTC, we suggest the model of display of the studied subject in spheres of the state activity.

The model encompasses the following: legal basics of the international cooperation, questions of a partner in the international cooperation selection, creation of criteria of efficiency of cooperation taking into account efficiency and safety, the organizational and administrative principles of the development realization at the level of parent enterprise of the project joint participants cooperation. The model of competitive price formation with allowance for target efficiency and advantages of production of cooperation compared to analogs relating to the tasks and other interests of the importer state is offered.

The presented material is based on the experience of concrete realization of the stated system approach of JSC NGO of Mechanical Engineering Military Industrial Complex, parent enterprise of cooperation, in the course of organization and handling of work in the field of military and technical cooperation at creation of rocketry, as well as the international cooperation in space area at creation of systems with radar supervision KA.

The authors of the paper under consideration study the issues of domestic high-tech enterprises competitive stability ensuring, and through consideration of their structural reorganization in particular.

Complex analysis of problems of large corporations, such as bulging executive personnel, decision making procedure foot-dragging, decline in efficiency of associations and their response to the market requests was carried out. As a result, the response to the market requests decline limits the growth potential. Under economic conditions of Russia, it leads to market sectors loss. Besides, under conditions of Russian developing market and a large quantity of imported products in the market, the enterprises of bulky structure would appear uncompetitive in the absence of qualitative restructuring.

The paper enlights the problems of Russian corporations and enterprises of micro-level management at the present stage of economic development, particularly, modern organizational structures of management (OSM) correspondence to the requirements of the new stage of economic development and states their improvement guidelines. It examines OSM relationships and existing market; analyzes the organizational and economic mechanism of corporate relationships with subsidiary enterprises, affiliates and organization departments, and offers recommendations for its improvement. The paper suggests an engineering model for improving the goals and policies on various levels of management in conjunction with their target trend.

The authors solved low-level problems concerning ensuring management of higher levels goals and policy by implementing the concept of «matrix-diagonal management structure». They suggest use it when developing OSM for scientific production associations (corporations), where production departments are vertical elements, while horizontal elements are executive personnel of systems production and scientific and technological support. A number of functions herewith, where the functions of subdivisionы and executives intermingle, are performed by diagonal elements.

Thus, the paper considers vital issues of an enterprise operation organization, their problems, wiсh hinder rate of economic growth at present, and keep on sticking for the most part to command-and-control methods of management, rather than economic. It leads to technical and economic slowdown.

This paper has scientific and practical value for radio industry enterprises aircraft engineering trend.

The paper presents the concept of estimation of labor potential effect on sustainable development of aerospace industry enterprise that includes three basic phases:

Phase 1: Airspace industry enterprise labor potential structure analysis.

Phase 2: Analysis and selection of aerospace industry enterprise activities that depend on specifics of aerospace enterprises and labor potential characteristics.

Phase 3: Evaluation of the enterprise cost with allowance for labor potential effect on its sustainable development.

Phase 1 starts with selection of sustainable development indicators of the enterprise, analysis of the enterprise specifics, methods and approaches to estimating the enterprise’s labor potential and factors affecting it.

At the first phase the optimum system of criteria for the particular enterprise according to the selected groups, as well as integrated estimated figures are formed.

At the second phase of the concept we reveal and select such performance indicators of the enterprise, which depend directly in a greater degree on labor potential and in a lesser degree on indirect and background factors. The authors selected two indicators: “product cost” and “enterprise cost”.

At the third phase the total cost of an enterprise is estimated within the framework of three approaches (extravagant, comparative and profitable) with allowance for sustainable development basic factor — labor potential.

As a result of structural transformations in the industry by the example of Scientific-Production Association “Saturn”, structural changes policy was formed for each of selected trends based on changes of labor potential quantitative and qualitative indicators: staff personnel, fixed-capital assets, raw material stock and taxes.

Reference [1] presents scientific and methodological basics for technical condition (TC) assessment of a guided aircraft air-to-air missile (GAAAM) guidance-system (GS). For this purpose new concepts of functional hardware capabilities and military potential of the missile were introduced, including the aggregate quality indicators, determinant of which is the missile terminal miss.

This article describes the tool and methodology to study and quantify the influence of GAAAM TC on the accuracy of its guidance, which can be used to implement the methodology described in Reference [1].

As a research tool the program-methodical system (PMS) was developed. Fundamentally, this system realizes complex mathematical models simulating targeting an aerial target (TAT).

With that, on the assumption of the volume and the depth of problems at hand, synthesis of the PMS and methodology for the impact assessment of the GAAAM TC on the accuracy of its targeting, involves the solution of a given problem in restricted sense and in a wide one.

The synthesis of this problem in a restricted sense is regarded as the accessible development tools (analytical tools) to study the required process or a specific object in a small area of varied parameters. In this form a similar PMS and the methodology were developed in [2]. Further it was tested in the thesis and several research papers.

In a wide sense, synthesis of the estimation method involves solving of a set of interrelated problems within the framework of the multifactor experiment (MFE) with the large dimensionality of initial data for objects research and processes investigated. It includes a wide range of initial simulation conditions and the simultaneous action of several different factors affecting the results of the MFE.

New modules were introduced in the developed PMS, including a module for the initial conditions setting for mathematical simulation of the guidance process based on approximate analytical dependencies. In addition, the additional cross-feedback connections were introduced and also a set of competing models were considered.

The outcome of the experiment caused by a number of other factors independent from each. Among them, we can highlight the following factors: types of used guided aircraft missiles (GAM); current technical condition of a used one, the type of a target and its flight performance (FP); the type and nature of enemy air target counter-effort to attacking missile; the GAM carrier type and its performance; the nature of air vehicle combat and the engagement nature with the target; the specific initial conditions of the GAM start-up onto the target.

In this regard, in PMS we applied five classes of modules to generate the initial MFE data, implemented by databases technologies: modules, determining technical condition of subsystems of GAMs; the modules that define the type of the target; the modules that define the GAM carrier type; the modules that form the initial conditions of combat use, the modules forming the end of the MFE simulation.

The PMS is based on a complex of mathematical models for GAM aiming at a target process, including a model of a target movement, the model of a GAM as an object of control (OC), as well as a set of models of the GAM guidance system as OC. The specified models form a guidance loop, closed through the equations of the relative motion of the missile and a target (RMMT). Complex mathematical models of the GAM aiming process at the targets is realized in the programming environment Borland C++.

Based on simulation results obtained with PMS for fixed conditions of combat use of GAM we obtained GAM and target motion trajectories; time depending functions of the change of the phase coordinates of the process of missiles homing at the target; 3D-functions of the quality indicator dependence from the two varied control parameter.

The features of a wing design of the modern medium-range passenger aircraft are considered. The main design deficiency of the classical single-fuselage scheme of the aircraft influencing the mass of a wing is the centrally located fuselage. The significant bending moment evolving thus in the wing root section is defining in weight calculations of its design. The configuration version of the twin-fuselage passenger airplane with a passenger capacity, identical with the single-fuselage airplane prototype, is offered. One of features of such configuration of the aircraft is presence of two instead of one fuselages with the pressurized cabins isolated from each other allowing single-level placement of passengers in salon at four chairs in the same row in each fuselage. This circumstance in addition to a flight safety and creation of more comfortable conditions for passengers allows to considerably increase unloading of a wing and reduce the bending moment in root section and the mass of its design. The power plant with three engines located on the wing of the plane creates additional benefits in comparison with the two-engine power plant at a possible engine failure. The four-leg fuselage undercarriage with two nose and two main legs also promotes increase of flights safety level in case of landing with a non-extended landing gear. The method of approximate calculation of a wing mass construction on the area of the bending moments diagram is offered. It is accepted, that the distributed mass loading of the wing design is function of the cross section area of wing bulkhead frame and normal stress is equal to the allowable stress. On the basis of the developed method the comparative estimate of a wing mass for the airplane-prototype and the twin-fuselage airplane scheme is made. Results of calculations confirmed efficiency of the wing unloading by fuselages removed from each other. The wing mass of the twin-fuselage plane appeared much less than a wing mass of the classical single-fuselage plane.

The article discusses the importance of the problem of reducing the life cycle of the propeller (propfans) because of foreign objects and elements of the weapons hits during combat operations connected with protected propeller blades (propfans) from exposure to particulate matter.

Classification of combat damaged items of military equipment caused by the impact of a weapon of the enemy and related factors is considered.

The paper describes the mechanism of formation and further damages evolution of propeller blades (propfan) made of composite material with the defeat of the elements of weapons.

The results of multiple damage simulated with the help of software CAE ANSYS / LS-DYNA are shown. These results clearly show that there is a change of the physical and mechanical properties of the material associated with the formation of bulk or surface hardening, which leads to the appearance of residual stresses, which is confirmed by experimental data. Also, there are the results of multiple injuries received in experimental studies of impact resistance with blade AV-112 propeller propulsion system of military-transport aircraft IL-112, which shows that the striking element is overcoming obstacles in the form of layers of fiberglass, increases the deformation zone that appears in a large area of the opposing barrier damage as size holes, and on the thickness of the material bundle and exfoliation area tape on the exit side.

The article shows the dependence of the area of separation against thickness of the bundle composite propeller blade resulting from an experiment and an example of the phase diagram characterizing the behavior fragments of the projectile in the form of a cube when struck by the blade in the composite spar at various angles of approach. The result is determined by the collision of impact velocity and the angle of the meeting.

In the article the authors point out the adequacy of the emergence and development of damage to the propeller blades (propfans) made from composite material with the defeat of the elements of weapons, confirmed by the results of calculations and experimental studies conducted in VUNTS Force VVA named after Professor N.E. Zhukovsky and Yu.A. Gagarin in the department Design of Aircraft Engines.

This work aims at searching similarity of RD170 engines family steady-flow operating procedures. We shall use generalizing functions describing interrelations of operating procedures parameters and valid for RD171M, RD180 and RD1971 for the formal description of the found similarity.

This work offers the reduction procedure of tests data to dimensionless form and defines the type and characteristics of the desired generalizing functions, i.e. dimensionless parameters dependencies on thrust level, which are common for all engines under consideration.

According to similarity theory, the propinquity of the equations describing behavior of dimensionless characteristics of the compared objects points to physical similarity of the processes described by these equations. Thus, obtaining the generalizing dependences reflects physical similarity of steady-flow operating procedures in structurally successive engines of RD170 family.

Based on the generalizing functions derived from the analysis of the results of fire tests by similarity theory methods, we developed an algorithm of parameters determination of multimode LPREs of RD170 family, allowing calculate their values in wide range of thrust variations according to known rated values of slow parameters (SP). The obtained results show that generalizing functions can find application for SP calculation in an infant state of LPRE of RD170 family design, when mathematical model is not correct enough, and the first experimental materials can form the basis for adequately predictive estimate.

The paper shows the possibility of using the results of presented work as references of normal functioning for solving the problems of function tests. The presented results of proposed method for firing test real time analysis within the space of which a throttle failure occured, confirmed the presence of such kind of failure.

Conventional commercial and military helicopters have flight speed limit of about 300-350 km/h. High-speed rotary wing aircrafts of the next generation require the increase of horizontal speed up to 450-500 km/h. Aerodynamic limitations, hindering flight speed increase of perspective high-speed helicopters, can be eliminated by reduction of the main rotor rotation frequency, while a helicopter builds up its high speed. This can be provided by implementation of regulated transmission with variable transmission ratio. The existing helicopter transmissions have constant transmission ratio. Thus, the development of regulated transmissions of anew type requires carrying out of exploratory study. The paper presents the results of the design of a new type of helicopter transmission with variable transmission ratio for perspective high-speed helicopters. It allows vary rotation frequency of the main rotor irrespective of engine and anti-torque, or pusher rotors rpm. The paper considers also the possibility of implementation and basic advantages and disadvantages of alternative types of regulated transmissions, such as, mechanical transmissions with stepped and stepless transmission ratio variation, as well as electromechanical versions of high-speed helicopters transmissions. It is shown that from the determining criterion point of view, i. e. minimum weigh and maximum efficiency of the transmission, the most effective and rational drive for a high-speed helicopter is a mechanical dual-mode transmission with stepped transmission ratio variation. The paper presents the description of a developed perspective stepless transmission scheme for a helicopter. The evaluation of dynamic forces acting during transmission transient modes with transmission ratio variation was carried out. Optimal scheme of the dual-mode mechanical transmission for perspective helicopters was developed. The important advantage of the developed scheme is the condition of safe operation, which consists in the fact that the coupling failure will not cause the braking of the force loop and loss of power at the main rotors. These design features provide high reliability of transmission operation and high functional characteristics with minimum mass increase in original structure of the helicopter main-rotor gearbox. This transmission mechanism is optimal solution for main rotors rotation frequency variation of perspective high-speed helicopters.

The article tackles the questions concerning peculiarities of vibroacoustic diagnostics of bearing assemblies for gas-turbine engine rotors technical condition. Possible reasons of bearing assemblies damaging and corresponding failure symptoms of individual defects, as well as procedures of their application are analysed.

The following options are possible in partucular:

• flight diagnostics (while processing post-flight information, without outputting warning the pilot at this stage);

• diagnostics during land approbation of the engine (aircraft-level) in operation or at the manufacturer engine test bench;

• during the parking, without engine activation, by manually turning the rotors or turbostarter cold scrolling (CS).

The last item is preferable due to minimum financial and working hour costs. For implementation of such methods manual turning of the rotor and a specialist fixing data or carrying out rotor CS are required.

In the course of diagnostics of operating gas-turbine engine aimed at researching possibilities of the technical microphone installed on the engine case a number of the methods connected with a high-frequency component of a signal spectrum of effective sound pressure was revealed. The studies consisted in searching of periodic and nearly periodic vibration processes resulting from bearing assemblies defect mounted on the engine and bearings tests bench.

High frequency periodic processes occur very often either separately from nearly periodic, or together with them. They appear at identical excitation of a vibration signal with each turn. These are frequency modulation of vibration signals from damaged bearing parts, as a rule, with rotor frequency.

Physical model of the process can be easily represented on the example of slider bearing operation. In the presence of negative factors, affecting the bearing clearance reduction, and occurring due to bearing capacity decrease or oil-film wedge punching, a mutual interference of wrinkles always presenting on sliding contacting surfaces takes place. It causes the excitation of oscillations of separate parts with frequencies equal to the product of number of interactions per one shaft rotation and rotor rotation frequency and natural frequencies of interacting parts.

The vibroacoustic diagnostics complex «FIANITE 3000» was developed to realize the data obtained during the studies.

It consists of a technical microphone (piezoelectric) with the restricted directional diagram, the DP-03 device, installed in inspection ports to inspect entrance edge of high-pressure turbine blades as well as electronic analysis and indication module FIANITE 3000. After DP-03 installation into inspection port and activation of electronic analysis and indication module the measuring process will go on automatically.

The developed system of vibroacoustic diagnostics is autonomous and protected from the main interference generated by subassembly of gas-turbine engine. The total evaluation of the system revealed its very effective noise immunity and serviceability for single and periodic checkouts.

The intensive studies are carried out on developing high-speed transportation vehicles, providing high level of environmental security, as well as higher energy efficiency of aircraft engines by means of fuel spraying and kerosene-air mixture combustion enhancement in aircraft engines. To improve the fuel spraying and fuel-air mixture combustion quality we suggest to use properly formed electrical fields in nozzle fuel supply contours. We considered for the first time the AC electric field of varying frequency impact on chemical composition of combustion products at the combustion-chamber outlet when using aviation kerosene TC-1 as a fuel. Moreover, we presented for the first time the experimental results of the of the AC electric field of varying frequency impact on the velocity of combustion products of air-fuel mixture.

Combustion products velocity measurement at the outlet of combustion chamber test model was carried out using the test bench at SGAU (Samara state Aviation Institute).

Electric field impact on the of combustion products of air-fuel mixture (kerosene) experimental research technique

The experimental velocity measurement of combustion products of an air-fuel mixture technique was developed at Samara State Aviation Institute and taken as a basis of the research on AC electric field impact on combustion products of kerosene-air mixtures velocities.

Employing measurement data, we calculated the gas superficial velocity and mass flow using well-known equations for gas-dynamic functions.

The result of the experiment allowed forming the file in Microsoft Access Data Base format with the possibility of export to Microsoft Excel.

The results of velocity profiles laser-optic measurements using 3D-LDA LAD-056C equipment.

The experimental studies were carried out at Samara State Aviation Institute with single-burner segment of the combustion chamber test model with serial double-contour nozzle of JSC “Klimov” for liquid fuel.

The swirler with blades angle set at φ = 72°10’, transition liner with outer cone diameter of 133 mm, square shape adapter with 180 mm side and basic variant of the mixer zone holes with all holes open were used . TS-1 kerosene was used as a fuel. Low-pressure compressed air was fed at ≤ 0.75 MPa. Solid tracer particles of CH-4 type for laser measurements were used.

When the AC electric field was applied along each diameter to the kerosene flow before its injection into the fuel nozzle the velocity changed to higher and lower values intermittently. Maximum relative decrease of kerosene-air mixture combustion products velocity at the outlet of combustion chamber herein was 2.45%, while maximum relative increase of kerosene-air mixture combustion products velocity at the outlet of combustion chamber in case of applying AC electric field to kerosene flow was 1.42%.

The design of modern power plants contains a significant number of coupling structural components of various shapes and geometries, made of materials differing in mechanical and heat-transfer properties. Most important task while creating a reliable power plant consists in correct calculation of thermal contact resistance occurring due to non-ideal contact of mating parts, and, as consequence, retraction and extension of heat flow lines to contact spots, as well as higher temperature gradient within the contact zone. All these factors reduce heat-conducting ability of the contact and cause different thermal expansion of the adjoining parts, leading to relative shifts, deflection and warpage of the parts.

The surface roughness is considered to be one of the key factors in solving thermal contact problems. Analysis of the literature on modeling and forecasting of contact thermal resistance shows that in practically all the works contain some analytical simplifications and assumptions concerning surface microrelief. In particular, irregularities were modeled as a variety of geometric shapes. The behavior of one pair of interacting irregularities was extrapolated to describe the behavior of a pair of interacting surfaces coated with irregularities. But if one takes into consideration the time when the suggestions of the CCC definition (60-70) were put forward, then we can say that the simulation of three-dimensional models was not carried out due to insufficient computer resources.

Despite this, there have been many achievements in the field of metrology and methods of numerical analysis.The optical measurement of surface features at the micro and macro level and the surface condition data storing in a digital form became possible. The numerical and finite-element modeling of contact problems with complex geometry, boundary conditions and material properties setting appeared.

Generally, to evaluate the temperature difference losses, the contact thermal resistance is introduced by different empirical formulas. But their diversity, incompletely given conditions of obtaining experimental data on which basis these relationships were obtained cast doubts on the correctness of the selection of a particular equation. While the differences in the absolute values of contact thermal resistance for the same conditions make it difficult to use them.

On the assumption of the formed problem, analysis of contact pairs of different materials was performed, and comparison was made with the known calculation dependencies.

Analysis of the known calculation dependencies to determine the CCC showed that the results obtained by rather accurate modeling of the ongoing process and analytical solutions differ, but they also have a number of matches. Nevertheless, the presented algorithm for calculating the CCC provides rather accurate values of temperature fields for almost all pairs of materials with minimal costs of the experiment. The use of this approach in engineering analysis allows reducing significantly the time of further testing and refinement of the product.

At the preliminary design stage of gas-turbine engines and high-temperature gas-turbine power plants one should use criterion dependencies to evaluate heat transfer coefficients from the gas at entry nozzle blade edges. Analysis of various criterion dependencies revealed that for the majority of correlations under consideration the degree of gas flow turbulence behind the combustion chamber was about (1 - 5)%, though for modern gas-turbine engine and high-temperature gas- turbine power plants Tu = (15-20)%. Information that behind the gathering main preceding the first stage nozzle set the degree of turbulence ε= 3 … 4% is not confirmed by the data obtained by Thole K. A. et al, Gandavarapu P., Ames F.E., Ames F. E., Nix A. C. at al, and in the area of maximum temperature field circumferential non-uniformity behind the combustion chamber. Thus, the paper by Thole K. A. et al shows that according to experimental results, verified by experiments and calculations with aircraft combustion cameras is gives 19%.

To compare various design procedures and identify the margins of their implementation the presented paper carries out comparison of criterion dependencies for the averaged and maximum local heat exchange at the entry blade edge with the results of 3D conjugated numerical calculation using ANSYS CFX and 2D calculations of turboprop engine nozzle blade with entry blade increased diameter.

The results of various techniques comparison revealed that H. Consigny and B. E. Richards averaged heat exchange criterion dependence should be used to evaluate the entry blade edge perimeter averaged heat exchange coefficients from gas. To evaluate maximum over entry blade edge perimeter heat exchange coefficients from gas, and in in the zone of combustion chamber maximum circumferential non- uniformity in particular (T^*_g = 2100 К and Tu ≈  20%), maximum heat exchange criterion dependence should be used. This dependency was obtained by the results of heat exchange while straight cylinder flow-around study carried out by Ekkert E. R. and Drake. As far as it is necessary at the preliminary design stage of turboprop engine nozzle blades to evaluate, in the first place, their high-temperature strength for the applied material, this design stage requires the use of criterion dependence . Position of maximum heat transfer coefficients over perimeter of the entry edge and their outstretch along its bumpy surface depends on many factors: gas backstreaming angle, gas turbulence intensity (Tu) behind the combustion chamber, the value of maximum gas temperature field circumferential non-uniformity behind the combustion chamber, and others.

Thus their location should be determined for each particular blade at the stage of 3D conjugated numerical calculations. The data, hereafter, on maximum local heat exchange coefficients on the entry edge external surface are verified experimentally.

Compressors and turbines of aviation gas turbine engines are the main components that determine the engine performances. Turbine blades are the most costly elements while their production. Rejection of an unfit turbine blade usually happens due to insignificant defects and consequently, their restoring is economically sound.

This article considers the process of geometry restoring (height) of aviation turboprop engine NK-12MP turbine starters (TS) blade wing. A comparative analysis of four types of restoration has been carried out: argon-arc surfacing with filler wire ХН60ВТ; soldering with solder powder VPr24 in ceramic forms; laser surfacing with filler wire ХН60ВТ; laser surfacing in powder bath with VPr24 solder.

The furnace soldering in ceramic forms technique consists in placing the blade in a specially prepared form repeating the ceramic blade profile with the necessary allowance for further machining, with further powder solder filling. Then the soldering process in a vacuum furnace begins.

The technique of laser welding in a powder bath consists in surfacing of the blade wing placed in a bath filled with facing powder. The blade was placed in the powder fill so that its wing and the surface of the powder fill were at the same level. The laser beam worked on the tangent of the blade wing and the powder.

Based on blade restoration techniques comparative analysis I concluded that the most effective technique in this case is pulse laser surfacing. Surfacing in powder bath herewith provides higher performance compared to laser surfacing with filler wire. Laser surfacing main differences from the classical techniques of turbine blades geometry restoring are revealed. The paper presents comparative analysis of the restored layers structure by electronic microscopy with elemental analysis of transverse sections of the samples. The degree of the effect of each technique on the blade basic material was revealed. It is found that the furnace soldering and laser surfacing techniques exert least effect on the strengthening phase γ’ of the cast alloy GC-6K. The elemental analysis revealed the presence of, presumably, the grid of complex intermetallic compounds, as well as tungsten and chromium carbides in the solder structure. Microhardness (Hv) of the recovered layers and various phases of powdered solder VPr24 was determined, and the CTLEM (coefficient of thermal linear expansion of material) involved in the restoration process was measured.

Spacecraft optimal control arrangement in exceptionally low-density Mars atmosphere, with allowance, in the first place, for the necessity to provide in the atmosphere the entry corridor of a required size,as well as effective deceleration of a spacecraft. Solution of the problem under discussion depends, in many ways, upon the proper selection of rational values of design and ballistic parameters of vehicles in descent mode and control methods in atmosphere.

The present work studies design-ballistic problems of a spacecraft descent in Mars atmosphere. It evaluates the characteristic trajectory parameters under various flight conditions of a spacecraft and analyses alternative control methods of a spacecraft. The obtained materials would contribute in many ways to substantiation of layout and rational technologies of spacecraft control during its letdown on Mars surface.

The article estimates physically realizable an entry corridor for a spacecraft with various aerodynamic quality values unchangeable in the course of the flight. It shows that with aerodynamic quality values decrease, the upper and lower boundaries of the entry corridor increase. The upper boundary of the atmosphere entry corridor of a spacecraft herewith is determined by overload maximum allowed value: the more the overload, the less the corridor upper boundary.

One of the ways of the spacecraft atmosphere entry corridor expansion is an effective quality management. It allows almost double the entry corridor width compared to implementation of spacecraft of a ballistic type moving with constant aerodynamic quality values.

An effective technique of final velocity extinguishing in atmosphere during the final leg of the flight of a spacecraft is introduction of soft landing system, which comes into action at the height of 5-9 km.

The advantage factor of the final velocity decreasing is reduced front surface loading reduction. Thus, reduction of the value from 350 kg/m² to 200 kg/m² leads to decrease of the values to 60-80%. In this conjunction, it seems necessary to carry out the studies of the dual circuit control of roll position and incidence angle of a spacecraft, which may provide significant reduction of the final velocity.

The article is devoted to theoretical study of the soft parachute-jet landing process. It proposes to use a long elastic linkage to suspend a load to the parachute.

Figura illustrates the operation of the parachute-jet landing system with elastic linkage. The engine thrust in the case of a hard suspension is 2.2 times greater than in the case of an elastic suspension.

The results of dimensions analysis for this problem reveal dimensionless groups, describing the process. A simplified mathematical model is built to describe the process of soft parachute-jet landing with an elastic linkage. As a result, of systematic numeric studies a rather simple interrelation between the dimensionless groups was found out. These obtained dependencies can be used for preliminary design.

Implementation of the elastic linkage allows significant reduction (several times) of soft-landing engines thrust. In its turn, it enables reduction of the following negative factors of soft-landing engines operation:

This enables you to reduce the following negative factors operation of the engines of soft landing:

• Significant vibroacoustic impact on a landing pad and airdropped object itself;

• Mechanical effect on a landing pad, which may cause its partial destruction;

• Considerable thermal effect of surrounding objects, which may cause a fire.

Application of the elastic linkage makes it possible to meet more fully the requirement on standardization of the existing parachute-jet soft landing systems with the possibility of continuous (not discrete) adjustment of system parameters to the given mass of an airdropped object.

The paper presents and analyzes the conditions and of a characteristics of the flight on modern aircraft with a crew of two pilots. The analysis of the flight is made in terms of flight safety. Examples of real events in the air, which led to the disaster due to the fault of the pilots, are described. Today, Airbus of A320 family performs the main air transport service, so the aircraft A320 control system is considered as basic for this analysis. The aircraft with a crew of two pilots control model in the form of man-machine system is built by converting the control system A320. Using mathematical tools of linear algebra the automatic and manual control of the aircraft structuring was carried out. Manual control of the aircraft is performed using SIDE STICKs by the captain and a co-pilot. More precise requirements to aircraft automatic control negative feedback were developed. The problem of the plane commander, which steer consists in the necessity to fly the aircraft in case of aircraft systems failure or incorrect actions of a co-pilot, was described mathematically. The model of the aircraft automatic control feedback transfer coefficients matrix was built in LabVIEW. Numeric and analytical studies of aircraft piloting in the form of the process of flight control were carried out. The definition of flight reliability was represented as a function of aviation equipment failure-free operation and error-free performance of the crew. The equations for flight reliability calculation under automated control and during active control of both pilots are presented. The regularity of reliability reduction due to aircraft automatic flight control system failure or non-participation of a co-pilot in controlling an aircraft is determined. Reliability calculation equations for operation in such conditions are presented. Detailed recommendations for flight reliability ensuring during crew preparation to the flight and flight operation are given.

This paper considers and algorithm of optimal landing trajectory shaping. Realization of this trajectory in automatic mode will maximally simplify the problem of transition to aircraft manual control.

Aircraft landing is considered to be the most laborious phase of the flight. The basis of the algorithm development consists in selection of automatically controlled trajectory as close as possible to the trajectories implemented by the pilot in the course of a manual landing, which will provide maximum convenience for the pilot in case of automatic landing failure. Thus, enhanced flight safety is provided.

The principle of the methodology consists in creation of formal specification of pilots actions during manual landing in the form of several optimization problems based on pilots actions in the course of manual landing. Thus, we consider the pilot as an optimal regulator, which performance criteria are sel ected according to the analysis of experimental data obtained earlier. As far as we consider manual control mode, the algorithm of control actions development should be made consistent both with emerged situation characteristics and with current characteristics of the pilot as well as the aircraft.

The paper analyses and formalizes flaring-out stage of the flight, which will provide landing safety. For this purpose, we realize on board the aircraft the flaring-out modeling algorithm with various options of throttle control and flaring-out altitude. Optimization herewith consists in selection of throttle control law and flaring- out altitude, wherein maximum regulator (a pilot, or ACS) error is tolerable.

Control performance of a pilot is considered in the course of studying of quasi-linear model which parameters are determined by recurrent identification in the process of flight realization.

We formalize glide-path capture in the form of the problem of optimal aircraft control, which criterion parameters are obtained fr om the experimental data analysis of successful variants of manual landing approaches.

The article considers the flight plan development language, used in software complex for automated spacecraft (SC) flight planning [1] to describe instruction structures [4] and prepare the initial flight plan data.

Design/methodology/approach

The language of the flight plan development is based on methods of structural, modular and object-oriented programming. [3] Like other high-level languages, the language of the flight plan development consists of characters, tokens (basic structure), expressions, operators and classes.

The problem of developing a new language stems from the need to develop a tool allowing describe commands used to plan the flight control and actions performed according to them.

The flight plan development language uses six types of data such as integer, character, logic, calendar date, time and n-tuple to store control commands parameters vector.

All calculations made are control commands execution time operations. This language defines variables for command time values storage. Each variable consists of type and name. The type determines the properties of a variable, and the name points for which command it is necessary to make calculation. Calculation of values is performed using the operations, which determine what time operations should be performed.

In addition to time operations, this language defines execution of conditional operations. Conditional operations are used to set the conditions for issuing control commands implemented during n a spacecraft communication with ground control station.

To provide repetitive calculations the language uses the cycle statement.

To describe more complex language constructions, used to compile a flight plan, «classes» were included. These Classes are used to describe the command structures. They represent a model carrying out an action executed by either SC, or a Ground Automated Control Complex. The elements that constitute such a structure can be both control commands and command structures themselves except the given one.

The language of the flight plan development was created for a formalized description of the command structures and the initial data used to compile a flight plan in an automated way. It allows describe the flight plan elements and structures more clearly, hiding the details of their realization. Its helps to provide the input data preparation for an automated spacecraft flight planning software complex.

Findings

This work results in creation of the flight plan development language for an unmanned spacecraft automated flight planning software complex. It aid in the description of structures and preparation of the input data used for compiling the flight plan in an automated way.

The developed language made it possible to formalize the description of the flight plan command structures, the initial data structure and simplify the way to describe them, while reducing the number of operators used in the description.

The created flight plan development language was implemented at the Mission Control Center (MCC) software as part of a complex of automatic spacecraft mission planning and has been used to prepare the initial data for the flight control of space crafts of scientific and socio-economic purpose.

Research limitations/implications

The flight plan development language can be applied during preparation of the flight plan initial data to control an automatic spacecraft of scientific and socio-economic purpose.

Originality/value

1. The flight plan development language has been created to describe the command structures and initial data preparation for flight plan made by an automated way with use of software complex of unmanned spacecraft automated flight planning. It was designed based on high level programming languages C ++ and C #.

2. The created flight plan development language allows structuring the initial data and to make it easy to read. Control commands are the basic elements of the flight plan. A simplified description of the flight plan structure allowed presenting a plan in the form of separate blocks and hiding from a human the details of its implementation using classes implemented to describe the command structures, which are based on the control commands.

3. The application of the flight plan development language for flight planning software complex of automated spacecraft has reduced the number of operators necessary to describe the initial data. This allowed reducing the time assigned for the preparation of the initial data and the number of errors introduced into the plan.

4. The flight plan development language as a part of flight planning software system was implemented in the MCC and was used for preparation of the initial data for an unmanned spacecraft of scientific and socio-economic purpose flight control.

The paper considers the problem of space application rockets' launching technological environmental impact reduction, particularly in the field of reducing the area allocated for the regions of rockets jettisoned parts impact areas. Nose fairing (NF) shatters are selected as a subject of research. The goal of the study is further development of impact areas reduction based on NF burning in dense atmosphere.

The paper sets out alternative techniques of NF shatters impact areas reduction, including representation of the suggested technique based on introducing thermite-igniting compound (TIC) to NF structure. As thermite-igniting mixtures, one can implement salts or metal oxides (KClO₃, KClO₄, CuO, etc.) mixed with one or several powder metals, such as magnesium powder, aluminum powder, titanium powder or their alloys. A binding substance, such as colloxylin is possible as well. The amount of mixture required to start the igniting process will depend on weight content of oxygen in the airflow, since combustion stability reaction just depends on its amount, and, hence, the required amount of heat emission as well.

The paper presents oxygen weight evaluation on the descent trajectory by the example of NF for the rocket carrier “Soyuz-2.1.v”. As a result, the time interval within which the combustion process should be realized. Evaluation of the developed technique adequacy was carried out.

The obtained results are the initial data for further development of NF burning in dense atmosphere.

Subject

The subject of this article is to analyze two aspects of application of the dual numbers for solving optimization problems of the multi revolution interorbital flight of the spacecraft with electric propulsion system.

Purpose

The purpose of this article is to demonstrate the possibility of using dual numbers in solving complicated optimization problems of the interorbital flight.

Methodology

The paper analyzed two aspects of application of the dual numbers to calculate the required derivatives during the solving optimization problems of the interorbital flight of the spacecraft with electric propulsion system. The use of dual numbers allows to determine the values of the derivatives with relative accuracy equals to the precision of function computation. The first aspect corresponds to the use of dual numbers with a single dual part together with the continuation method for calculating the elements of the sensitivity matrix of the system of nonlinear equations corresponding to the boundary value problem of the Pontryagin maximum principle. The second aspect is the use of dual numbers with vector dual part to calculate the right-hand sides of the system of differential equations, which describes the optimal process. By virtue of the canonical formalism of the maximum principle it is necessary to calculate the optimal Hamiltonian in the dual representation. This approach is used for solving the optimization problem of the interorbital flight when the model of the spacecraft motion takes into account various disturbances.

Results

The results of two types optimization problems of the interorbital flight of the spacecraft with electric propulsion system between the initial elliptical orbit and geostationary orbit are presented in this paper. The first type of problem corresponds to the motion model of the spacecraft in the central gravitational field under the influence of the reactive acceleration. For the second type the model takes into account the effect of the Moon and Sun attraction, and disturbance caused by the Earths gravitational field noncentrality. The functional that corresponds to the minimum mass of required fuel quantity is considered for both problems. Solution of the first type was obtained by using the continuation method and approximation of elements of the sensitivity matrix using dual numbers with a single dual part. For the problem of the second type, dual numbers with vector dual part is often used when we determining the right- hand sides of differential equations of optimal motion of the spacecraft. The nonlinear boundary value problem in this case was solved by a hybrid Powell algorithm.

Practical implications

Overall, the methodological approach of the using dual numbers for automatic differentiation outlined in this paper can be used for solving any optimization problems, where model of the object motion is described in a complicated manner.

Conclusions

This paper shows, that the using of the device of the dual numbers for computing required derivatives can efficiently solve complicated optimization problems of the interorbital flight of the spacecraft with electric propulsion system.

One of possible approaches to determine kinematic error is functioning modelling in MSC. Adams, which allows investigate static and dynamic characteristics of machinery and mechanisms.

The following errors affect the kinematic error of harmonic gear drive with rolling bodies:

• diameter error of rolling bodies;

• wave former diameter error;

• a rolling body mounting separator aperture sizes error;

• profile of a rigid wheel manufacturing error;

• rigid wheel teeth uniform distribution error;

• wave former axis of rotation position error.

The article presents the results of wave gear with rolling bodies mathematical modeling in MSC.Adams at nominal deviation of sizes of certain parts of the gear.

The kinematic error of the gear at a constant rotation frequency was determined, and its harmonic content was analyzed. Dependencies of kinematic error components on the values of the gear links sizes deviations, such as wave former disk diameter and rolling bodies diameter, is revealed.

In accordance with the results of the studies, we can make the following conclusions:

1. wave former disk and rolling bodies sizing errors lead to occurrence of gear kinematic error, with one static and one harmonic component, which frequency is determined by the rotation speed of the wave former and the number of rolling bodies of the harmonic gear;

2. the backlash of the harmonic gear with rolling bodies linearly depends on an errors in sizes a wave former and rolling bodies;

3. the cyclic component of wave gear with rolling bodies kinematic error is approximately of a linear character depending on an error in sizes of a generator wave former and rolling bodies.

At present, the requirements placed on power actuators include increase of load capacity, positioning accuracy, reliability, durability, efficiency, impact resistance, and a number of other parameters. The required speeds and accelerations of the output element of the drives are increasing as well. Drives should be easily mounted on the object, replaceable and adjustable. Their servicing should be simple, and their control should be reliable and easily programmable. For most products, especially of the space industry, it is necessary to reduce the weight of the drives and maintenance facilities. The environmental requirements placed on the drives are constantly toughening.

The analysis of the requirements placed on the drives for special aggregates reveals that they should provide:

· Fast response and playback accuracy of a desired motion law;

· High precision of moving;

· Synchronous operation of two or more drives;

· Greater amount (range) of the output element speed control;

· Higher soft ride of large mass products at low (micro) speeds;

· Minimal and stable energy losses values.

In addition to the abovementioned, a power actuator should have a simple technological design, low weight and size, and high operating reliability. Recently, the demand for the planetary roller-type actuators (PRTA) as the actuating mechanism (AM), especially for the military and aviation equipment, increased significantly in our country. A number of companies are trying to master serial production of PRTA, which requires to address the issues of PRTA analysis validation and design, production technology of the actuator parts and their assembly, as well as high-precision metrological control of PRTA parts critical dimensions, using the appropriate equipment.

PRTA has a number of advantages:

· High load capacity;

· The possibility to provide light feed at high load;

· High reliability;

· Low sensitivity to temperature variations;

· Ease of mantling and dismantling of the transmission;

· Ability to work with high rotation speeds;

· High efficiency;

· Low metal consumption.

Based on the conducted studies of mathematical models and well-known techniques [3, 4], a technique for a prospective PRTA AM structure and parameters justification was developed.

To solve such problems, the experts of FSUE «CENKI» — «DB «Motor», MADI and N. E. Bauman MSTU proposed to use the planetary roller-type transmission as the actuating mechanism.

Theoretical and experimental studies under the scientific supervision of Professor A.N. Sova and candidate of technical sciences A.V. Sizanov are conducted on the manufactured friction planetary roller-type transmission 48×12, and the control system on a specialized workbench for transmission running test is prepared.

Research is conducted at the branch office of FSUE «Center for exploitation of space ground-based infrastructure», — «Design Bureau «Motor» using the CNC milling machine Hardinge Bridgeport GX1000. A measuring device is sel ected based on the desired accuracy of transmission. Digital measuring head Mahr Extramess 2000 with permissible error G = 0.6 μ m was used in the study.

Conclusion

Thus, implementation of scientific and methodological approach and the conducted research allow us:

· Take into account the longevity reserves of contact fatigue damage, in-service wear, and quality of the lubricating layer;

· Determine the optimal value of the actuator mass to ensure a pre-set speed without overshoot;

· Reduce deviation of the experimental results fr om theoretical results;

· Develop scientific and methodological support of the results of the machine-aided design of the actuator with high dynamic performance, reliability and durability implemented with CAD/CAM/CAE system.

At present, planetary roller-type transmission is the most promising device that converts rotational motion into linear.

High-tension polymeric composite materials (HTPCM) are widely used in aviation industry due to their high strength-density ratio and modulus of rigidity; good shock-absorbing capacity; corrosion resistance and low thermal-expansion coefficient. HTPCM tooling, however, is associated with a number of difficulties. Its low thermal conductivity affects the heat balance: the main part of heat, about 90%, is concentrated in the cutting area and at the drill tool, whereas up to 70% of heat leaves the cutting area with chip scrap while tooling metals. This high temperature causes partial melting of polymeric matrix, tempering of tools and processed surface. Composites anisotropy of properties leads to differences in chip formation along and perpendicular to the reinforcing fibers and their high hardness causes increased wear of the drilling tool. Delaminations, caused by axial force and torque of the cutting process, are formed in the places of drilling tool entering and exit.

The purpose of this work is to study the impact of the structural and geometric parameters of drilling tools on the quality of the processed holes, productivity, dynamics and thermal physics of the high strength polymeric composite material machining — titanium foil reinforced glass fiber plastics.

Extra loaded polymeric composite constructions implement titanium foil reinforcing packs, allocated between the layers of reinforcing glass- or carbon ribbon.To provide required accuracy in size of apertures macro- and micro-geometry, elimination of tempering and delaminations of the material one has to use sharp-ground tools and carry out processing in stages. This work studied drilling process of fiber-glass laminate at various schemes of high-speed steel drill tools sharpening.

The work presents the results of processing rates optimization for the studied tools geometries; dependence of unevenness tallness parameters; the character and value of delamination in the places of drilling tool entering and exit, depending on the tools type and processing rates, as well as qualitative image of chip formation. Using the developed computerized real-time measuring system for power characteristics of drilling process, we found blunting criterion and efficient life of the cutting tools under study. The pyrometric registration data on average temperature in cutting area allowed perform finite elements modeling of thermal processes occurring while multilayer plastic, reveal the sources and directions of the thermal flows, and, finally, develop recommendations for technology improvements of apertures processing in glass fiber beams of main and steering rotors.

Among mathematical models describing various processes associated with manufacturing engineering, aerospace technology and aviation, there are models representing a system of ordinary differential equations and a system of nonlinear algebraic or transcendental equations, i.e. a system of differential-algebraic equations (DAE). Such problems often arise in applied mathematics and mechanics. Some hydrodynamic processes described by the DAE systems contain a small parameter (viscosity). Problems of this type describe the phenomenon of creep. The process of creep of the material in the first approximation can be modeled by DAE systems discussed in this paper.

The basic methods of solving boundary value problems for such systems are methods of collocation and shooting methods. With shooting method, a boundary value problem is reduced to some initial value problem. However, this method is applicable only in the case when the original problem is correct. We suggest to apply the best parameterization for regularization of this problem.

The paper considers the system of nonlinear differential-algebraic equations. It is shown that the best parameterization of the boundary value problem for a singularly perturbed differential-algebraic equations significantly improves the computational algorithm of the shooting method.

The numerical solution of the problem was obtained using the method of solution continuation with respect to parameter and the best parameterization. The boundary value problem is reduced to an initial value problem for differential-algebraic equations. We selected shooting method as a numerical solution.

By using these methods of solution all solutions of the boundary value problem were obtained regardless of the choice of initial values.

According to the results, following conclusions were made:

• the method of solution continuation with respect to parameter and the best parameterization can be used for solving singularly perturbed boundary value differential-algebraic problems;

• the method of solution continuation with respect to parameter and the best parameterization allows to find all solutions of the boundary value problem for nonlinear differential-algebraic equations.

Numerical studies of this work show that the parameterization of the boundary value problem for nonlinear differential-algebraic equations, proposed in this paper significantly improves the computational algorithm of the shooting method and allows to find all solutions of the boundary value problem. Thus, in this paper we propose a numerical method, which allows solve the applied problems related to technology of machine building, rocket and space technology, aviation.

Synchronous machines with rare earth permanent magnets are widely used in aviation and rocket technology as motors for the electric actuators and generators of electrical power systems. These machines are promising for use as the main aircraft generators with a high level of electrification in the DC power supply system of increased voltage 270 V. They are characterized by a large variety of designs, most important of which are versions with radial or tangential magnetized permanent magnet and nonmagnetic or bimetallic holder. Constantly increasing demands to the mass-energy performance of aviation synchronous machines with rare earth permanent magnets, to improve the quality of their designs, the development time and cost reduction are met by continuous improvement of design methods and design techniques based on the calculation and analysis of electromagnetic fields in their active zones.

Rational combination for electromagnetic fields in active zones analysis is traditional methods based on the models with lumped parameters together with circuit theory and computer technologies of numerical finite elements analysis based on the models with distributed parameters and field theory. Traditional methods and created on their basis by way of setting relatively simple and easily observed mathematical interrelations between input geometry and windings parameters with output energy parameters new design methodologies allow quickly and solve the problem of initial active zone geometry of an electric machine and parameters of its windings selection with an accuracy sufficient for engineering analysis. We used numerical methods for electromagnetic fields verified analysis and optimization, where parameters obtained with traditional methods are set as the starting point.

Traditional methods of synchronous machines design and machines with rare earth permanent magnets in particular are based on a number of assumptions and system of calculation coefficients linking the parameters of a real distributed active zone model with lumped parameters of electrical and magnetic circuits. For example, excitation field of synchronous machines is commonly characterized by a system of calculation coefficients κ_f, κ_ф, α_{δ и} κ_Β, determined based on the analysis of excitation winding two-dimensional magnetostatic field or permanent magnets in the smooth working gap brought form the armature side with non-saturated armature and inductor magnetic cores.Saturation consideration is carried out, as a rule, by separate introduction of various correction factors.

An effective method for determining the calculation coefficients of synchronous machines is the method of harmonic analysis of magnetic fields in active zones of electromechanical converters. The article presents new analytical solutions obtained for the magnetic field in the active zone of synchronous machines with rare earth permanent magnets by harmonic analysis. The problem of calculating the magnetic field generated by the permanent magnets of sector form with a constant direction and magnitude of magnetization, located between he two cylindrical ferromagnetic areas with infinite permeability was solved in polar coordinates system. The resulting solution allows determine the excitation magnetic fields and calculate coefficients of synchronous machines with permanent magnets of sector type, as well as prismatic and circular forms with their allocation on the rotor and fastening them with nonmagnetic holder, as well as reversed structure. Also, the problem of the magnetic fields calculation of excitation and armature reaction of synchronous machine with permanent magnets and a bimetallic holder with alternating magnetic and nonmagnetic areas was solved. The solutions presented in the article were checked by comparing the calculation results with the results of numerical finite element analysis. The results mismatch herewith does not exceed 1%.

The spacecraft power supply system as well as ground-based power supply system provides transmission and redistribution of electric energy from power sources to consumers. One of the major requirements claimed to power supply systems is to keep electric energy at specified level in the course of one or more primary energy sources failure.

Currently, two types of supplying electric power to consumers exist: centralized and decentralized. The structure of the centralized power supply system has the Central Switch Gear (SSG) that collects electric energy from all primary and secondary power sources.

The main advantage of such systems is the stable feeding of consumers due to required power extraction from all electric power sources. On the other hand, the centralized electric power system has relatively low power grid mass and size figures, as well as comparatively low reliability due to the presence of single distribution gear.

The prospective trend, with the presence of a number of primary and secondary power sources in particular, is implementation of decentralized electrical power supply system, having several small SSGs, which potentially increases reliability and scalability.

The article introduced the concept of an agent and multi-agent power supply system that provide the operation, storage of information in the database, as well as exchange of information with other agents that allows implementation of the decentralized power supply system.

Consumers, sources and energy storage units are the main elements of the system. We suggest designing a separate agent for each element, which will provide — functionality, information storage in database, as well as information exchange with other agents to provide their effective operation.

We identified the following agents: a load agent (LA), an environment agent (ENVA), a photoelectric converter agent (PHCA), a fuel cell agent (FCA), a battery agent (BA), a simulation agent (SA) and a database agent (DBA).

The article also presents a power circuit of a static electric energy converter based on modern RB-IGBT transistors and method for synchronizing the output AC voltage with other power sources. It allows improve conversion efficiency of the energy generated by solar cells (this is one of the agents of the decentralized systems), reduces the mass and size parameters of the converter.

The article is devoted to the issues of competitive stability of Russian aircraft engine-building enterprises amid rapidly changing market demand. It stresses the necessity to develop a whole number of measures and the aggregate of product and process innovations aimed at quality improvement of output products; expansion of productive activity diversification level; carrying out the restructuring of companies of aircraft engine-building sector; entering world market as suppliers of 2^nd-3^rd level components; implementation of non-traditional finance-and-economy tools and schemes to provide an enterprises entering the world market as a supplier of competitive aircraft engines.

The pressing problem for national aircraft industry is formation of globally competitive world-class engine- building sector. It is planned to increase budgetary allocations volume for aircraft engine-building at the expense of Federal budget from 1 billion 448 million rubles in 2015 to 3 billion 431 million rubles in 2025. With such a significant financing increase of the sector, the compulsive issue of competitiveness of the enterprises is challenged. This concept means time-phased competitiveness of the aircraft engine-building enterprises. We need to preserve gained competitive advantage as longer as possible, as well as raise a level of intra industrial cooperation. This will require the increase of economic responsibility of all parties for the results of their common labor and achievement of common goals. It will also require intensification of all manufacturing processes and enhancement of individual role of all structural divisions of an enterprise.

The aircraft equipment life cycle is a significant and special component for the competitiveness analysis due to its scope of production and time parameters of implementation. Thus, organization of aircraft equipment competitive production should take these specifics concerning life cycle with allowance for upgrading possibilities, capital and restoring repair with allowance for assigned resource for components, assemblies and aircraft in whole into account. Mandatory elements also are maintenance and technical support as well as service maintenance of products when in use.

Development and advancement of new products of the enterprise of the aircraft engine-building sector should be carried out with emphasis on international cooperation, allowing develop global chains of deliveries of their own and create strategic alliances with other aircraft building market participants, including efforts within the frameworks of core innovation territorial clusters' activities. Global servicing network that will help the companies of this sector to transfer to the full life cycle of the engines management is going to be created.

The aerospace sector features a highly considerable specifics and types spectrum of management problems. Among these problems the presence of management decisions optimization problems is mandatory. Correspondingly, management goals in aviation and space-rocket sectors are diversified, heterogeneous in dimension, conflict and knowingly irreducible to a single utility measure. Thus, in all cases the multicriteriality occurs.

A certain conceptual idea of a group of optimization criteria notations (interpretations) diversification has the right to existence.

In various fields of mathematical modeling identification and algorithm presentation significant progress can be observed, which may lead to creation of tools allowing newly interpret management situations and find new solutions for management optimization problem, based on conceptually new mathematical methods, namely numerically or even of closed form.

The requirements on presentation include:

— Correspondence to the basic fundamentals of scientific cognition;

— Universality with relation to number and conceptual definition of optimization criterions, as well as procedures applied for their evaluation;

— Keeping the sequence of optimization criteria positioning;

— Keeping the dimensionality of original optimization criterions;

— Keeping mutual priority of original optimization criterions;

— Ensuring representation additivity (componentwise addition feasibility);

— Using such structures of scientific theories within which frameworks the inception of solving optimization problems methods is possible;

— Vector optimization criterion applicability for correct scalarization procedure.

The group of management decisions optimization criterions in aerospace sector can be represented (interpreted) by implementation of the following conceptual schemes:

— Conceptual scheme based on the theory of sets. In this case, the original group of optimization criteria is presented in the form of a set K, such that the original optimization criteria are declared disjoint singleton subsets of this set:

;

— Conceptual scheme based on linear algebra. In this case, the group of optimization criteria is presented in the form of a vector , such that the original optimization criteria are declared components of vector criterion:

;

— Conceptual scheme based on the classical algebra. In this case, the group of optimization criteria is represented as a function K, defined as n-th degree polynomial of conditional argument A without constant term, such that:

;

— Conceptual scheme based on the theory of functions of complex variables. In this case, the group of optimization criteria is presented in the form of functions K of complex variables with parameters , defined as a complex variable without the real part, such that:

.

It is obvious, that all abovementioned interpretations provide for the tend to the extremum, which is also a component (maximum or minimum).

Home producers of civil aerotechnics are not able to enter foreign market and almost ousted from the home market due to severe competition with foreign companies. One of the reasons for the present situation stems from the lack of methodology of civil aerotechnics competitiveness evaluation, which allows objectively determine level of competitiveness and the ways of its increasing.

Research objective — developing of aircraft competitiveness evaluation model, which allows for all basic factors of cost and non-cost character.

Employed methods — analysis methods, deduction and forecasting.

According to our research, the existing approaches are not always effective.

The proposed model searches for competitors, evaluates their advantages and disadvantages, makes a comparison of level of costs and takes into consideration passengers’ judgements. Thus, we evaluate the aircraft competitiveness at the market.

The article motivates the necessity of comparing only the aircrafts with equal levels of both customer service and airworthiness certificates.

The paper proves the necessity of airliners operating costs with minimum cost of flight per one passenger criterion, since only with this criterion gives the possibility to consider all operating costs, ICAO restrictions, differences between aircrafts passenger capacity and course speed.

One cannot use the current price level for calculation of aircraft life span operating costs. We recommend use predictive values for POL price indices; cockpit and cabin crews labor costs, currency courses, airport charges rates etc. Thus, we obtain minimum dynamic flight cost criterion.

With reference to passengers’ preferences, the author means their requirements of flight comfort, personal area, air conditioning quality, low noise level and high cabin pressure.

Aircraft building companies can evaluate the competitiveness of their product using the described technique. The proposed model should be useful for airline executives, while selecting an airliner.

Since 1994 diplomatic and economic relations between Russia and Brazil differ in the positive dynamics of political contacts on all levels. October 2013 marks 185 years since the establishment of the Russian-Brazilian diplomatic relations.

In a few last years and especially after introduction of sanctions directed to level deceleration of Russian economic integration into international arena, the relation between Russia and Brazil develops with the high rates. Most changes take place in the field of external economic collaboration, especially in the field of science (in particular, space programs), energy and finances.

In opinion of Goldman Sachs, to the 2050 the summary economies of BRICS countries will exceed the total size of economies of the richest countries of the World (The Group of Seven).

The set dialogue determines external economic partnership between Brazil and Russia. A primary goal is to diversify bilateral trade with the aim to lead its volume up to 10-11 milliard dollars in 2020.

At the present time, Brazil continues the development of operational satellites SCD-1 and SCD-2, which were first launched in 1993 and in 1998 year respectively.

An important area of cooperation is a joint modernization of the Brazilian VLS-1 rocket (so-called program “Southern Cross”), which is the Brazilian entry ticket to the club of space-rocket countries.

However, nowadays in Brazil only the use of space developments in the field of telecommunication brings the financial return.

Another area of cooperation between Brazil and Russia is the nuclear power industry. Today, Brazil has the need for energy.

In the framework of the VII meeting of the Russian-Brazilian high-level commission on September 6, 2015 a Memorandum of Understanding between the State Corporation “Rosatom” and Nuclebras Equipamentos Pesados SA (NUCLEP), the leading company of the nuclear industry in Brazil, providing services in the field of heavy engineering, has been signed.

New opportunities for enhanced cooperation between Russia and Brazil are opened through the Business Council of BRICS and there are already examples of successful cooperation projects in the energy sector, in particular, the localization of production of power equipment in Brazil.

The growing interest in renewable energy sources associated with the steady growth of energy consumption, as well as increased emissions of greenhouse gases into the atmosphere.

The main task of Russia is search of ways to increase the commerce and, as consequence, expansions of investment cooperation. The level of investments has noticeably grown for the last some years, possible risks here became significantly less, but also the time of reception of the first profit has significantly increased. It is connected with the enough limited home market of consumption in Brazil, where most part of the population lives below a level of poverty and only buys products of the first indispensability. Thus, manufacture orientation to the export trade is more preferentially.

Conclusions

Cooperation between Brazil and Russia develops fast paces. The important directions of cooperation between two countries are such high technology branches, as space branch and atomic engineering. During cooperation Brazil aspires to take the experience, saved up in Russia, for progress of own technologies. Also, the interoperability between two states in area of a science and formation, tourism and other branches actively develops. The Fundamental importance for both countries has regional cooperation.

This article is devoted to the realization of aircraft engines competitive recovery in the framework of the State program of development of aviation industry in Russia until 2025. It is proved that aviation engine-building belongs to the number of high-tech and knowledge-based industries, which development affects a number of vital areas of life and, in whole, the stability of the national economy.

One of the important trends of economic efficiency and Russian aircraft engines competitiveness increase is to ensure their reliability on the level of foreign competitors. The urgency of the system of aftersales service development and repair time reduction through maintenance workflow optimization using progressive norms and standards. The problem of optimal division of labor should be addressed comprehensively, taking into account economic, organizational-technical and psychophysiological factors, which are closely related.

Thus, the article describes the method of division of labor optimization in the organization of aircraft engines maintenance of using models of integer linear programming and network planning. The paper presents the results of the implementation of a method of optimizing the division of labor occupied with assembly operations for aircraft engine maintenance.

Hence, the use of these innovations will allow provide uniform loading at the workplaces, reduce the number of workers and the complexity of final assembly.

Active realization of State programs of industrial and economic development of Russia in modern conditions of managing increases the need for ensuring more mobility of its population. In conditions of Russia, avionics can provide the most effective transportation of passengers with minimum time losses, and implementation of light helicopters in particular.

For steady development of business in the sphere of rendering helicopter services in the territory of the Moscow region, it is necessary to provide the high safety level of air transportation. Safety characterizes the degree of security from impact of risks, internal in the first place. Safety is a basic significant meaning for risk management and ensuring the greatest possible degree of security of social systems from economic and anthropogenic impacts.

The development of helicopter transport depends on a solution of the problem of development and implementation of modern techniques and technologies of risk management. In economically developed countries, the assessment of airline activities efficiency puts on the first place the security of its activities, rather than the profit made. The risk is considered herein as a market grade, which identifies to a great extent both the position, and competitiveness of the rendered service and its producer.

The subject for the study is a risk management system while operation of commercial helicopters. As a result, the concept of risk management based on safety notion is suggested. The airline internal risk assessment tool is developed and tested, and the measures aimed at risk management system formation are elaborated on.

The study of aerodynamics influence on the motion of an artificial cluster-satellite of two bodies is stipulated by the expansion of large extent orbital tether systems field of application. For such large space systems, aerodynamic effects are of a great significance. We have investigated the behavior of limit cycle for an equation of relative motion orbital cluster of two bodies considering impact of gravitational effect, aerodynamic pressure, airborne gradient and dissipative factors, depending on the growth of aerodynamic parameter a. We use the equations of interconnected motion with allowance for the forces of gravitational gradient and aerodynamic factors. We study the effect of aerodynamic parameters on the behavior of limit cycles. To carry out the research we implemented well-known methods of nonlinear mechanics: of Lagrange equations of the first kind method, phase plane method; points mapping method; theory of motion stability methods.

The main qualitative effect of the rotational motion of the satellite in an elliptical orbit is the possibility motion chaotization. The atmosphere creates especially strongly effects the beginning of chaotization due to the exponential change of its density in an elliptical orbit. Even relatively small eccentricities can enable strong chaotization.

For eccentricity values e = 0.001 and small values of the aerodynamic parameter a limit cycles are absent. The increase of aerodynamic parameter a leads to the emergence of limit cycles. For values of a fr om 4 to 45 there a limit cycle exists. The value of (( to which all points on the phase portrait roll down drops from 29.6 to 14.5. With a further increase of the parameter a (a > 45) lim it cycles disappear.

Design of a side-stick control (SSC) of the aircraft is an important task, which is focused on improving the efficiency and convenience of pilot operation. In addition, application of SSC allows reduce the total weight of the control system.

The paper carries out the analysis of present-day SSCs, describes the effect of added mass inherent to any SSC. It also shows the disadvantages associated with control priority characteristic to passive SSCs, and shows that active SSCs have a number of advantages over passive ones due to permanent synchronous operation.

The paper presents the results of morphological analysis and synthesis of SSC development. It is shown that construction of SSC gimbal joint using two electromechanical drives placed over skew axes allows reducing weight and size figures, and implementation of double-reduction harmonic gear with rolling elements makes it possible to design hybrid control channels. It is found that application of double-reduction harmonic gear with rolling elements for power mini drives with output stage operating as differential mechanism allowing summing up motor and mechanical guide motions is expedient.

The paper reveals that introduction of additional mechanical guide for manual control increases the level of pilots information awareness as well as flight safety level in emergencies.

The paper contains the results of design, mathematical modeling and the appearance of the prototype. The model is based on dynamics of a drive system, taking into account dynamics of permanent magnet synchronous motors, robustness of the design and reducing gear, moments of inertia, and nonlinearities, such as reducing gear backlash and dry friction.

The authors developed a number of algorithms of SSC operation in different modes and described the method of setting the control priority. It is shown that the presence of force sensors improves the functionality of the device.

The paper outlines the structure and operation of the aircraft control system when SSC is in use, and shows that transfer to manual control in case of power-off is possible.

The presented results demonstrate that using of hybrid control channels increases the flight safety level.

Aerodynamic characteristics of large elongation straight elastic wing with turn winglets are analyzed. The discrete vortex method is used for determination of the aerodynamic load distribution. Mathematical model of the wing structure is developed using the bay method as a finite element method with enlarged elements of a thin-walled beam subjected to bending, transverse shear and torsion. These methods are used to solve coupled aeroelasticity problem. Aerodynamic load distribution for elastic wing and its divergence velocity are determined. Dependence of the aerodynamic characteristics from the winglet turning about longitudinal axis for some of its locations are analyzed.

Aerodynamic symmetric and antisymmetric load distribution taking into account the turning of the winglets is considered in this paper. The winglets can duplicate and even execute the functions of the wing regular mechanization elements as ailerons. The use of winglets can be particularly efficient for small aircraft with large elongation elastic wings.

For determining the aerodynamic load the wing is divided on small finite panels with one joined vortex and two end free vortexes of constant unknown circulations. Flow non-separation conditions are satisfied in one point of the panel.

Using bay-method the wing is divided to bays by cross sections. The bay stiffness matrixes are determined using the theory of thin-walled beams. Displacements, angles of rotations, and twisting in the joints of the bays are considered as generalized coordinates. The total stiffness matrix is constructed by standard procedure of finite element assembling.

As an example aerodynamic load distribution and divergence velocity are calculated for straight wing with turn winglets.

The proposed technique makes it possible to estimate the efficiency of the turn winglet control by aerodynamic and aeroelastic characteristics of wings. Calculations executed for typical structure of straight wing with large elongation show that using of that winglets may be efficient tool for wing aerodynamic characteristics control. The winglets can act as ailerons to control the antisymmetric motion of aircraft.

Using of the turn winglets evidently is most reasonable for light aircraft with rather compliant wings of large elongation.

The paper presents the method for economic effect estimation when using new computer hardware (compact super computer) to calculate aerodynamic performance of unmanned aerial vehicles (UAV). It vividly illustrates that computer-aided experiments (calculations) reduce drastically the amount of testing carried out in wind tunnel. By expert estimation of foreign specialists (USA, Great Britain, France, Germany, Israel), depending on flight mode and UAV configurations for which the aerodynamic performance is determined, the wind tunnel tests cost reduction is five times less and even more. In this case, the reduction of tube testing herewith does not lead to loss of volume of information, necessary for high-quality design.

The paper presents the method for calculation of manufacturing process economic component to estimate the aggregate of required expenses and economic effect analysis of a research work on computation cluster design for numerical study of aerodynamic performance. As practice revealed, this method can be successfully applied for economic appraisal of the abovementioned research work conducted in 2013-2015.

As the results of the selected method for economic effect calculation revealed, the expenditures connected with acquisition of a new compact super computer as well as carrying out mathematical modeling are rather significant. But they become justified compared to the number of blowing-downs in a wind tunnel and related explicit costs.

For example, with six testing blow-downs in wind tunnel the expenses are equivalent to the cost of new equipment. Moreover, when the number of estimated six blow-downs is exceeded, the results of comparison show that compact super computer zone of effectiveness comes.

Requirements placed on modern perspective helicopters envisage not only vertical take-off, landing and hovering, but long-range cruise speed as well.

Having the same engine installation, a rotary-wing can achieve high flight speeds due to optimal use of the main rotor, additional fixed wing and propulsion plant.

Creating the lifting force, the wing unloads the main rotor, preventing the retreating blade stalling. Propulsion plant sets optimum fuselage attack angle when it does not create negative lifting force and has minimum drag.

Implementation of jet engines to create helicopter propulsion force is not effective since they are not used during hovering. Thus, in case of using turboprops as cruise engines not the entire power of main engines is realized at high flight speeds. This drawback can be rectified by using one or more extra propellers, driven by common transmission. In case of single rotor helicopter propellers can be used for reactive torque compensation at low flight speeds. At high speeds reactive torque is compensated by vertical tail.

The alternate method of propulsive force creation at high speeds consists in implementation of tilting the tail ""rotor. In this case, all the power of main turboshaft engines will be used at hover, as well as at high speeds, distributing between main and tail rotors.

The main problem of rotary-wing design is selection of the wing and propulsion prop parameters. This problem was solved by parametric analysis using Mil Moscow Helicopter Plant flight simulator. Engine power and main rotor parameters were fixed, and wing area and propulsion prop thrust varied.

For a single rotor helicopter with two VK-2500M turboshaft engines and take off weight of 11500 kg, additional 15 m2 wing should be used to reach maximum cruise speed 385 km/h. The main rotor autorotation is not an optimal way to achieve the highest aircraft performance characteristics. The main propeller should compensate fuselage drag. It needs 2500 h. p. for this purpose. The main rotor, with this, produces 70% of required lifting force for a horizontal flight and consumes 1500 h. p.

The article considers design techniques for reliability and safety upgrade of space-rocket systems (SRS) operation.

The authors have suggested the reliability demonstration technique of the systems when measuring operational integrity parameters, as well as solved the problem of SRS reliability demonstration control under amount of testing limitations. Information content of tests has been enhanced. The paper defines the areas of effective and bearable values of the parameters under study. Failure probability was evaluated for each parameter through coefficients of variation of effective and bearable parameters values and safety margin. Interval estimation was calculated to provide guaranteed result. Average margin value was evaluated using operational integrity values sampling. For the accepted level of confidence, lower and upper boundaries of reliability were calculated. The sampling can be arranged using testing results carried out in the course of experimental development. Thus, amounts of redundancy introduction provides SRS reliability requirements for the limited amount of testing. It is very important in relation to time limit.

We suggested systems reliability demonstration technique during extra heavy testing, and solved the problem of SRS operational integrity control during contingency rating. Availability of redundancy of system parameters characterizing its operational integrity is suggested. Loading factor was determined on account of product operation in both effective and bearable modes. After extra heavy testing the lower boundary of reliability for the accepted level of confidence was determined using tables. Thus, the introduction of amounts of redundancy provides the SRS heavy testing reliability requirements. It is of prime importance concerning one-shot products.

The paper suggests an SRS failure occurrence evaluation technique, and solves the problem of a bad product acceptance probability.

The technique of SRS validity control parameters optimization is offered. The problem of optimum tests number determining with allowance for tests expenses and inspection error damage is solved.

The algorithms using the offered techniques, which are illustrated with modeled examples, are presented.

The received results can be useful for the design offices workers, research institutes, scientific and production enterprises during developing, as well as operation and maintenance of highly reliable SRS.

The paper presents the results of the study of SPT-100PM and SPT-140PM stationary plasma thrusters models characteristics. These models operated with different schemes of discharge feeding and high discharge voltages. The abovementioned models are the SPT laboratory models with external accelerating channel diameters of 100 mm and 140 mm, respectively, modified for magnetic shunt allocation inside the discharge chamber. They had reduced width of the accelerating channel and magnetic system optimized for operation with high discharge voltages.

Tests of these models had shown that due to the abovementioned modifications we managed to reduce the negative impact on thruster performance of mass flow reduction while discharge voltage is increased, to provide moderate level of discharge power, as well as accelerating channel broadening due to discharge chamber walls erosion during continuous operation. It is also shown that:

• the best performance level could be obtained with two stage structure of the discharge feeding scheme when magnetic shunt potential is negatively shifted relative to the anode potential;

• it is possible to obtain an anode thrust efficiency of the SPT 140PM (calculated not accounting for the cathode mass flow rate) model within the range of 0,550,60 and its «anode» specific impulse up to 4000 s, when the sum of the discharge voltages on both stages is up to 1400V and discharge power does not exceed 5 kW.

The efficiency enhancement of aircraft engines through quality improvement of fuel spraying and fuel-air mixture combustion is the question of the day. The results of experimental studies of properly induced electric field, using electrical facility affecting aviation kerosene (EFAK) at the fuel nozzle input, as well as chemical composition of kerosene-air mixture combustion gases were obtained.

Based on the carried out analysis it is possible to draw the following inferences:

• cascade connection of the EFAK flow part with the fuel nozzle without supplying voltage to EFAK electrodes changes relative mass content share of the corresponding chemical element of combustion gases at all exhaust points. Moreover, depending on the type of a chemical element and exhaust point, these changes can either decrease, or increase in relation to the results of previous measurements;

• on EFAK electrodes energizing the content of С_хН_y, CO, CO₂, H₂ in combustion gases changes significantly compared to Х_uо. Moreover, depending on exhaust point these changes can bear the signs of + or — Thus, for the point of probing of combustion gases placed at 20mm from the left edge of the horizontal cross section diameter of combustion gases flow the decrease of C_xH_y content by 22.191%, CO by 17.5367% and CO₂ by 9.53% with simultaneous increase of H₂ content by 3.7313% occurs. For the probing points at 60 and 100 mm from the left edge of the combustion gases flow cross-section the increase of С_хН_y content by 10.5263 % and 8.3832 % correspondingly takes place. The content of CO₂ increases by 6.5638% and 2.8604%, while the content of CO increased by 14.927% and 3.9484% correspondingly. Moreover, the content of H₂ decreases at these probing points by 3.7313 and 0.7692% correspondingly;

• such non-unique changes of chemical elements content at these three probing points in the course of applying voltage to EFAK electrodes can be explained by possible non-uniformity kerosene TC-1 drops flow distribution over the sections during combustion in the combustion chamber.

At the majority of enterprises running aircrafts a considerable part of documentation is prepared in paper copies: maintenance schedule, units data, service bulletins, updating information, technical servicing manuals, spare parts and units lists, their certificates, logbooks, aircraft operating time information etc. With such immense information volumes in the form of bulky paper copies the process of search, selection, filing and adding new data represents a time-consuming operations. Solving the above-mentioned problems requires development of a software and its practical application in the system of technical servicing of aircrafts. As a whole, implementation of a software in the field of provision of technical servicing will significantly enhance the efficiency of aircraft utilization. The paper presents the developed software allowing carry out in full the recordkeeping of all aircraft failures in electronic format and issue scheduled reports on detected failures and their processing in accordance with domestic regulatory documents. Its distinctive feature consists in realization of 3D model of an aircraft in mobile application synchronizes with identical model allocated in the server. It is very easy to mark on it the damages revealed during the inspection. A photo of this damage together with its characteristic is transferred to the server, and logged in trouble record log.

The developed software — is a mobile Android application encompassing all main steps of aircraft failures control: from failures journalizing to accompanying documents generation. In general, implementation of such software in the area of servicing provision will significantly increase the efficiency of aircraft utilization and servicing, which will contribute positively to flight safety level.

The developed software was realized in ASYPOLA system. It is consisted of two subsystems: Android_part and Personal_Computer_part solving the following problems: automated failures recordkeeping, gathering, storing and filing of information, substituting paper copies.

The author chose power electromechanical actuator drives based on planetary roller-screw gear (PRSG) of improved accuracy and reliability of operation of the special equipment for integration installation of large-sized with and large mass products with a rocket carrier of a super-heavy class as the object of study.

The technique includes the preparation of the initial data and the basic requirements to the actuator, an electromechanical drive circuit design. It also considers selection of materials and heat treatment of the main parts of the mechanism, the selection of lubricant, the design of the actuator, a preliminary calculation of the transmission efficiency, as well as power and kinematic analysis of the transmission. It allows calculate power losses in the mechanism, and adjust the calculation (if necessary), creation of a model to determine the main characteristics, checking compliance of transmission with specifications, comparison of the options and selection of the optimal one, the creation of the final model, test strength calculations, preparation of design and technical documentation.

As long as power electromechanical drive is a vital unit with high requirements to accuracy of manufacturing of certain elements, the solid-state modeling with adjustment at each stage is necessary. The solution of such problems is made difficult without the use of through-design CAD / CAM / CAE systems.

To ensure the required speed and accuracy of the output displacement of the electromechanical actuator we use the actuator with high reduction and efficiency. This is planetary roller-screw with a number of advantages over conventional transmissions:

• increased load capacity due to power transfer by several streams with load distribution between rollers and increase of contact lines in helical links total length;

• allows improve reproduction accuracy of a given law of motion of the mechanizm;

• reduces weight and size of the mechanism and provides a high overall efficiency due to the large reduction of the stage;

• performs frame and supports load reduction, due to the distribution of forces in the planetary gear;

• allows complete elimination of the gap in the transmission;

• provides high smoothness of motion, allowing high accuracy to perform a given output displacement law depending on the input signal;

• provides high smoothness of movement when moving at micro speed;

• exhibits small friction losses and high durability;

• operation capability in a wide range of temperatures and in vacuum;

• operates at high speeds and acceleration;

• has no impact on the environment;

• compactness.

The basic elements of the transmission are as follows: screw nut, threaded rollers, toothed crowns, separator, locking rings. Both screw and nut can be the key element in this transmission. Screws and nuts have a thread of triangular shape with an apex angle of 90(. Rollers have a special thread which profile in axial section has the shape of convex arcs of a certain radius, due to which a point contact is provided in conjunction with the roller screw and nut. The design of such transmission allows for different angles of helix of screws and nuts.

Justification and development of scientific and methodological apparatus, and creation with its application of electromechanical drive with planetary roller-screw transmission for space nosecone tilters of a space-rocket carrier of super-heavy class allows improving the process of preparing a space-rocket carrier for launch. Improving the process of space-rocket carrier preparation for launch means increasing accuracy in reproduction of the ground tests program with regard to the laws of nosecone motion and allows reduce the possibility of emergency situations due to incorrect operation of the installed equipment and machinery of a space nosecone at the boost phase.

The paper considers a flat contact problem on an elastic half-plane with structurally non-uniform surface wear-out by the rigid wear-free punch. While stating the problem we suppose that elastic characteristics of the material are structurally tolerant and identical in all points of the body, and parameters characterizing wear- resistant properties of the material depend on the value of linear wear. We also suppose predetermined the main vector and the main moment of force applied to the rigid punch.

To solve the contact wear-out problem under consideration we implemented variational approach. As a result, we obtained the problem formulation in tensions in terms of the system of quasivariational inequality of evolutionary type and first-order differential equation. For the problem time sampling, Euler explicit difference scheme was used. Thus, to determine contact pressure an elliptic inequity or extremum problem equal to it should be solved on each time step. We performed the problems quantization over spatial coordinates using the space of integrated fundamental solutions of the Flamant on the effect of normal concentrated force over the surface of elastic half-plane. We used boundary element approach to plot the surface of integrated fundamental solutions. Elements with a uniform distribution of contact pressure were used. As a result of sampling we formulated the problem of quadratic programming with restrictions in the form of equalities. The linear transformation of variables, allowing simplification of the restrictions was suggested. For numerical solution of the problem we used a variant of conjugate gradients method taking into account specificity of restrictions.

The calculations carried out showed that structural heterogeneity over the depth of the wear-out surface significantly influences the nature of the interface wear-in process.

The developed computational algorithm can be implemented for study of the efficiency of various technologies for superficial hardening by the method of computing experiment.

The paper considers a problem of identification control synthesis for mobile robot. The problem emerges in cases when it is necessary to solve the problem of the control system synthesis for an object, which mathematical model is completely unknown. Examples for the occurrence of such a problem while robot guidance may bepresented as follows: the presence of an object described by a highly complex mathematical model; problems for which the use of the model in full is not necessary, or mathematical model of the robot is not known completely due to external, environment, as well as in case of partial loss of functioning.

In this paper, we use a numerical method of the network operator to solve the problem of identification control synthesis for mobile robot. This method allows determining the structure and parameters of mathematical expressions in the form of integer matricesusing evolutionary algorithms. The main advantage of this principle consists in using small variations of the basic solution. The principle lies in the fact that the researcher defines one basic solution, and the evolutionary algorithm searches for the optimal solution on the given set of basic solution variations. This approach allows building a set of possible solutions, where the majority of the functions satisfy the condition of the control goal. The principle of small variations reduces as well the searching range, limited to the vicinity of the basic solution. In the process of searching, basic solution is periodically substituted by the best current solution. The network operator method was effectively used to solve various problems of synthesis and identification.

The problem of identification control system synthesis consists of two tasks: identification of mathematical model of a controlled object and a control system synthesis for the obtained model. Initially, we apply the network operator method to solve the problem identification, in which, the quality criterion is a trajectory deviation from the experimental data. The experiments were performed on the real object, using the control, close to the required for control system synthesis problem. At the second stage, we solve the control system synthesis problem by the network operator method as well. We use another network operator of different dimensionality and with different quality criteria. We realize the network operator obtained as consequence of synthesis in the control unit of the real object and verify the results of identification control problem solution.

The results of conducted experiments displayed satisfactory quality of identification and control synthesis problem solution.

Schemes of so-called pulse-modulated multilevel voltage inverters with output sine wave, three level (3L NPC) in particular, widely spread thus far, both in home and foreign publications on power electronics. Such inverters are implemented in the devices operating at high switching frequency and requiring high efficiency (low switching losses) and high quality of output energy. Such devices require also low harmonic content and corresponding to it noise emission with acceptable mass and dimensions, dynamic and reliability parameters of output filters, such as uninterruptible power supplies (UPS), solar batteries inverters, onboard frequency converters, etc.

Three level inverter cell (totem pole) of three-phase off-line voltage inverters with output sine wave represents traditionally a half-bridge pulse modulator based on input two-capacitor structure with grounded center tap (for three phase variant in particular). It also consists of unidirectional four-switch transistor structure with free-wheel and grounding diodes, and output LC circuit. The output LC filter (Lf-Cf) acts as demodulator (low-pass filter). Depending on its parameters relationship it provides various dynamic time lag and degree of robustness (slope ratio) of inverter output volt-ampere characteristic near its operating point. Conduction losses in three-level inverter (3L NPC) are slightly higher, than are those in two-level one. However, three-level inverter provides significant switching losses reduction and allows decrease total power dissipation by 40%, which is especially impressing at high switching rates. SEMIKRON company manufactures SEMITOP & MiniSKiiP dedicated modules based on IGBT intended for invertors design within power range up to 100 kVA. Their structure provides all traditional protection circuits: overvoltage, overcurrent, short circuit and overheating. Moreover, in contains circuits for power switches turn-off in case of through-currents occurrence and active voltage limiting of the power switches.

As an essential shortcoming of methods and means of protection, implemented by the company, we should mention traditionally straightforward approach of foreign companies, according to which they proceed from invariable structure of the power stage (without any attempt to its revising). They pay attention only to control scheme and forming switching trajectories by control signals parameters. Thus, the designers fully neglect theoretically substantiated and tested in practice means, such as soft switching and overcurrent and overload protection, as well as switching heat losses and noise emission reduction. These means are: inductive and capacitance non-dissipating (energy) damping and snubber circuits (DSC), which guarantee switches zero-current turn-on and zero-voltage turn-off. Moreover, there are no attempts to exclude circuits for possible through-currents and inverse diode over currents, disregarding the possibility of their impact of powerful EMI of lightning on driver circuits (regardless of logic circuits of guaranteed pauses etc.).

Another, no less substantial drawbacks, typical to this approach, are the following established traditional approaches to equipment design. They are: a) implementation of input energy-consuming capacitive filters based on electrolytic capacitors with rather low reliability factor (thermal stability, durability, life cycle); b) the lack of inverse current gain-phase adjustment circuits (both for idle and regenerative current), running in a rectifier mode and a corresponding power factor correction, significantly reduced due to AC filtering inductor (ballast) and switching rectifying action; c) implementation of step-down invertor and rectifying modes only, significantly narrowing functional features of conversion (the input-output voltage ratio range); d) low output electricity quality of output power (output voltage form and parameters stability). All above mentioned features do not allow implement well known inverting and rectifying converter circuits to realize universal module design and module-scalable architecture.

With reference of all above said the authors of the paper think that the attempts to implement the circuit design means leading to increasing production and operational effectiveness of so-called bidirectional inverting and rectifying converters (BIRC) do not exhaust themselves. The paper represents the examples of some possible approaches to the matter by means of radical BIRC power circuits upgrade, protected by RF priority.

The original approaches to radical modernization of the traditional three-level sinusoidal voltage inverter proposed in the article are applicable for the design of unitized modules of bidirectional (convertible) inverting and rectifying (or else, rectifying and inverting) converters (BIRC or CRIC) with demodulating and storing reactors, power factor correctors and non-dissipating damping and snubber circuits (for soft switching aimed at reliability and efficiency increase and noise emission reduction).

These circuits can be recommended for the design of unitized modules meant for the synthesis of multi-functional (multi-phase in particular) switched mode converters (MSMC) with functions of direct BIRCs, CRICs, convertible frequency converters (CFC) and switched mode converters, regulated sine currents and voltages inverters (RSC/VI), uninterruptible power sources (UPS) and other types of converters for electrical complexes of module-scalable architecture, which provides high production, assembling and servicing processability, as well as high energy saving, reliability, weight and size and price effectiveness, as well as high quality of power and EMC factors.

This article seems to be interesting to a wide circle of designers of power electronics, especially in the field of aircraft on-board energy supplies, and for designers of fully electrified aircrafts in particular.

Purpose

The flying vehicle on-board cable network is subjected to electromagnetic interference due to its length. Electromagnetic interference immunity and susceptibility of on-board systems strongly depends on the on-board cables shielding efficiency. It is one of the main characteristic of flying vehicle on-board cables. When conducted current induced by electromagnetic interference flows over the external surface of shielding cover, the shielding efficiency is evaluated on the basis of well-known frequency response of coupling impedance. Coupling impedance of shielded cable can be calculated based on theoretical models or found experimentally with co-axial test setup. Coupling impedance models of shielded cable differ due to various approaches of describing the physical phenomenon of the in-phase voltage inducing in inner conductors of the cable, when current flows through external surface of shielding cover. Frequency responses of coupling impedance models differ from each other, which leads to the deference of frequency responses of external shielding efficiency, calculated on their basis.

The goal of this work consists in studying of coupling impedance models of on-board cables for flying vehicles, their divergence at the specified frequency range and comparison with experimental characteristics obtained using co-axial test setup.

Design/methodology/approach

The theoretical part is based on classical electrodynamics and electrical engineering. Elements of the complex variable theory and operational calculus were used as well.

Findings

The paper presents the results of theoretical and experimental coupling impedance and shielding efficiency frequency response of on-board cables for flying vehicles plotted in MATLAB. It also studies the differences between theoretical and experimental coupling impedance and shielding efficiency frequency response at specified frequency range. It is marked that shielding efficiency frequency response is valid for case of conducted current flows through the cable shielding cover in the absence of external electromagnetic field. In computing shielding efficiency the current through shielding cover must be of conductive origin rather than induced by the field of an external noise, i. e. it can be, for example, a reverse receiver-to-source current in case of non-symmetric connection. If a current through external shielding cover is induced by electromagnetic field effecting the on-board cable, the calculation of shielding efficiency based on coupling impedance frequency response can lead to significant erreos.

Current, which flows on cables shielding cover, should be conducted electromagnetic disturbance when shielding efficiency is calculated. For example, this current could be common-mode return current from the receiver to the source. Shielding efficiency calculation based on coupling impedance frequency characteristic can lead to significant errors in case the conducted current induced by electromagnetic field.

Completeness of the reseaerch

The presented paper is a part of the research on calculation of electromagnetic interferences induced in on-board cables of flying vehicles.

The field H(r,z) of axially magnetized torus in the vicinity of its axis of symmetry Z has been numerically studied. It is shown that the profile of the field allows construct a simple mathematical model to determine ponderomotive forces in the torus — orb system. Unloaded (in the absence of external forces f = 0) soft magnetic orb in the field of the torus has three equilibrium positions near H_max: in the center of the torus, over the torus and under the torus (areas C, A and B, respectively) — Fig. 1

Stable position of the orb under the load f is realized under condition of dH(r,z)/dz < 0. Integral equation for orb magnetizing can not be solved by such model, since it assumes a simplified calculation of the orb polarization in the field of the torus with surface charges averaging over two hemispheres of the orb, and solution of the forces affecting magnetic charges of these hemispheres.

The model allows further simplification by introducing into consideration of effective magnetic charges located in the center of gravity of the hemispheres charges on the Z-axis. The forces acting on the point charges are calculated. Effective charges provide the charge neutrality of the orb. In this manner, the model is reduced to one-dimensional model. The model simplifies the calculations without loss of accuracy. Radial r-component of the field of the torus is much smaller than Z-component. —component of the field H(r,z) is relatively linear in the computation domain. It serves as a basis for the introduction of the one-dimensional model.

In this paper, the calculations were made for the torus of the size of 9,7 × 4 × 1,25 cm. The magnetization of the torus is set to M_z = 915 emu/cm3. The saturation magnetization of the orb is set to M_z = 1620 emu/cm3.

The diameter of the orb is 2R = 1,15 cm, the gravity weight is equal to 6,121·103 dyne. The displacements of the orb under its own weight relative to the unloaded orb: in the area A — ∆A = 0,353 cm, in the area B — ∆B = 0,741 cm, where ∆A < ∆B .

The maximum values of ponderomotive forces, at which the position the of the orb is steady, in area A — F_Amax = 2,072·104 dyne, in area B — F_Bmax = 6,846· 103 dyne. The ratio of the maximum forces in the areas A and B is equal to: F_Amax^/F_Bmax = 3,027. When |f | > F_max, an abrupt reposition of the orb occurs: fr om area A to area C, and from areas C and B to (- ∞).

Levitation ranges in the area A — ∆A = 1.164 cm, in the area B — ∆B = 1.217 cm, wh ere ∆A < ∆B Displacement of the unloaded orb relative to the maximum value of the field is equal to: +10-3 cm in the area A and −103 cm in the area B. The area C is not considered in this paper.

Numerical calculations explain the experimentally observed significant difference in the displacement of the orb in the field of gravity and maximal forces F_Amax and F_Bmax above and under the torus. When the diameter of the orb is greater than it reaches the area of the sign reversal of the torus field, and the maximum value of the force is not defined in the model.

The experiment showed that the ratio of maximum calculated force F_max to the force measured in the area A is equal to 1,37. It is obvious that the accuracy of the model increases for the samples of smaller diameter.

Taking into account the given above data it can be concluded that the model can be implemented for semi- quantitative evaluation of the power characteristics of the system torus-orb.

Operation algorithm of traditional multi-phase inverter structuresfor multi-level output voltage formation assumes consecutive summing of voltages from individual cells along the period of the function. According to the proposed algorithm, sinusoidal reference signal is divided into equal voltage levels in accordance with the number of inverter cells so that the output voltage approaches sinusoidal waveform.Then we determine the intersection points of voltage levels with the sinusoidal signal and time intervals for each level.The increment valueof voltage levels is the same. It depends upon the number of cells, and only the durations of each level vary. As aresult we obtain approximated sinusoidal signal. By increasing the number of cells, the output voltage approaches the form of the reference sinusoidal signal.

Each time interval of a voltage level operation is divided into equal intervals, according to the number of inverter cells.Then, using the proposed algorithm of power switches commutation, we obtain the total voltage of each level by summing voltages of all cells both in serial and in parallel.

This algorithm allows uniform distribution of the loads of all cells, which, in its turn, allows using similar cells for multilevel inverter.Other well-known schemes for formation of an approximate sinusoidal output signal does not provide an equal cells load.

The proposed algorithm of operation of power switches of cells is clearly structured, and allows commutation scheme scalingfor any number of cells.Theoretically, the maximum number of cells of a multi-phase inverter depends on the minimum switching time of a single power switch, comprising a cell of an inverter.

The study of simulation results show that there are practically no low-order harmonics when the number of cells equals to six.

As a rule, high frequency switched mode power supplies (SMPS) are used for reactive magnetron sputtering [1 3]. The point is that such pulse powering allows reducing the probability of electric arcs occurrence and increases stability of reactive magnetron sputtering process. One of the main parts of SMPS for magnetron sputtering systems (MSS) is a transistor former of output pulses. It determines the power source type (unipolar, bipolar, HIPIMS) and its operating efficiency.

As is well known, in contrast to continuous input current mode, switching losses, peculiar to switched mode limit maximum switching frequency, and subject transistors to pulse overloads decreasing the reliability of a pulse former. If soft commutation of transistors is provided, zero-voltage or zero-current switching will allow reduce dynamic losses in transistors and increase reliability of a pulse former.

Reference [4] presents modular SMPSm based on bipolar pulse former (BPF), where soft switching was provided by using snubbers, limiting the current or voltage growth rate during transistors switching.

However, to ensure correct operation of snubbers as a part of BPF their adjustment is necessary, with allowance for real parameters of components in use and BPF operating algorithm. The present work is devoted to solving this problem.

In the course of this work, we developed a computer model of BPF in Swicher CAD/LTspice, based on standard SPICE models of power electronic components. With the help of BPF model, we calculated parameters of snubber elements, necessary to ensure soft switching conditions and energy recovery stored in snubber elements.

In addition, the energy efficiency of BPF with snubbers at the maximum frequency of the output pulses forming of 75 kHz was calculated. The calculation results showed that the use of the snubbers allows reduce significantly switching losses of transistors. Due to energy recovery, stored in snubber components, we managed to reduce the total power losses of BPF by 25 % and increase pulse former efficiency by 0.6 %.

Switching losses reduction of transistors provides increased ВPF reliability, which together with higher efficiency offers a prospect of a further increase of the output pulses frequency and improved stability of reactive magnetron sputtering process.

The paper considers some specifics of small innovative enterprises foundation in Russian gas-turbine engineering.

The percentage of small innovation enterprises (SIE) that pertain to the space-rocket hardware and gas turbine equipment in Russia is currently low. This can be explained by the fact, that commercialization of the results of scientific research in such a science-consuming industry as gas turbine engineering, where prospects are of no guarantee, presents high risks.

The author considered the example of a small innovation enterprise foundation in Russian gas turbine engineering with participation of a large gas-turbine engineering company and venture capital. SIE 1 has for its object development and commercialization one of the advanced competences in the sphere of parts design, development of industrial technologies and short-run production for gas-turbine engines.

At a project startup stage in gas-turbine engineering, Russian practice is characterized by common mistakes, which may turn into a real threat of risk situations occurrence in the sequel. These threats involve insufficient presentation; incompleteness of the products and services; the absence of unbiased price-quality relationship; low efficiency of labor caused by its wrong organization; ignoring legal questions of the case; irresponsiveness for the project executing; input barriers and volume of investment underestimation.

The set of methods for economic risk assessment, suggested by the author, allows get reliable information about the risk in tome. Such information is necessary for risk management, and can be recommended to assess investment qualities of investment objects during foundation of innovative small economic environment of Russian gas-turbine engineering companies.

Since 2005 the Russian — Latin American commodity turnover increased twice and by the end of the 2015 first quarter totaled 17.8 billion USD. Latin America is the only region of the world, for which the increase in Russian exports is observed.

The goal is to consider the problems that arise in the process of cooperation between Russia and the LAC countries in the region, factors that have a direct impact on changes in the external economic relations of today, as well as to determine the prospects of their development.

The subject of this study is external economic relations, which determine the state of the level of interaction between the states, regional groups and other parties of the world economy, gives the opportunity to defend the interests of other States in the field of economy, as well as identify some of the theoretical aspects of international relations.

This defines the range of problems underlain the presented research and its structure:

• Analysis of the disputed and controversial issues of modern Russian development strategy of foreign economic relations in order to develop a clear position of Russia in relations with the countries of LAC;

• Reasoning and consideration of external economic relations, against the background of the political situation in the international arena;

• Using logic relationship between objective and subjective factors in the development of foreign economic relations.

High practical relevance of the theme is stipulated by the great interest to the issues of foreign economic cooperation between Russia and the LAC countries, which do not hold a strong position at international import and export markets yet. Analysis of these issues has both theoretical and practical significance.

The scientific novelty of this study is as follows:

• A generalized analysis of the economic situation of the LAC region in terms of attracting foreign investments and defining the main directions of development of foreign trade of Latin American countries with the Russian Federation in the future is carried out;

• The range of the most promising trade and economic partner of Russia in the of LAC countries is defined.

The field of application of the results may contribute to the development of a number of measures at the state or municipal structures level for the development of foreign economic strategy with the involvement of practically poorly known regions of LAC.

To sustain at present the leadership in the field of national aircraft engineering in the interests of design and serial production of new types of aircrafts the corporate development plan, including long-term ones, should contain such items as creation and development of modern design department and production, which operate using digital technologies, the newest software, up-to-date test-bench and laboratory base and new flight test cluster. The strategic development way of an enterprise is formalized in the form of a complex long-term planning of technical development with due regard for the existing material, technological, information and human resources. The current annual development planning (the nearest project) and a three-year planning (medium-term project) are formed based on this long-term planning.

Nowadays an urgent problem of defense manufacturing divisions modernization and their provision with skilled personnel revealed itself in full scale. Furthermore, it is noted that the scientific and technical reserve of Russian enterprises of the defense industry complex is almost exhausted. Therefore, the complex strategic development project should include long-term planning to establish the scientific and technical reserve and preparation of skilled personnel to maintain and improve technical, technological and intellectual resources of an enterprise. Human resources development as well as specialists training are among the main the priorities of JSC Tupolev.

Andrey Nickolaevich Tupolev, the founder of the company Tupolev, has shown a great example of the strategic thinking and realization of advanced ideas. He relied on the existing scientific school as one of the elements of the strategic planning. The enterprise hasn’t lost this accumulated scientific and technical experience, but nowadays its not sufficient. Only new knowledge, skills and administrative qualities can ensure the success.

It is necessary to know the state-of-the-art management and marketing tools. The necessity to preserve the succession of the majority of structural and engineering solutions when transferring from one aircraft type to the other improved types of aircrafts should form the basis of PJSC Tupolev and other modern aviation and production organizations strategies. It is also necessary to cooperate with fundamental and applied sciences organizations, which form the main scientific reserve of the industry and participate in the tests of aircraft samples under development.

To implement strategic development projects of the enterprise we suggest form a group (or department) of highly skilled professionals. The main missions of such group will be acquisition and analysis of the existing scientific and technical information, participation in the carried out research works within an enterprise, organization and systematization of development and implementation process of industry technologies, long- term development planning related to the aviation science and technology.

Strategic development plans implementation should allow the design engineering center enter a new scientific and technical level, enable to design, develop and realize tests of the modern competitive aircrafts.

The main indicator of an airline carrier competitiveness is air tickets low sale price as long as safety and comfort standards are observed. Accordingly, aircraft manufacturers face an urgent problem of producing an aircraft least-cost in operation. The existing operating costs assessment criteria do not allow for environment versatility, ICAO restrictions, various passenger capacity, and other factors. It leads to unacceptable errors when assessing passenger plane operating costs.

The paper pursues an objective of developing a criterion, considering prime operating costs, ICAO restrictions, various passenger capacity and course speed, as well as environment inconstancy.

According to ETOPS regulations the flights of the planes equipped with two gas-turbine engines must be performed in such a way that in case of the failure of one of the engines the flight duration with one operating engine to an emergency airfield from any flight way point under conditions of zero wind and speed set for the flight with only one operable engine would be 60 minutes for non-ETOPS aircrafts, and up to 270 minutes for Boeing 777.

Thus, the only true criterion, which is able to assess a long-distance passenger plane operating costs objectively, is flight cost value per one passenger.

The cost of one ton of the fuel, labour cost of flight technicians, flight crew together with flight attendants, currency rate, airdrome fees etc. are not constant over the total lifespan of an aircraft. Thus, there is a danger that the flight cost value indexes would display unacceptable error when comparing aircrafts with differences in performance and economic characteristics.

Thus, to assess operational costs objectively we recommend use dynamic cost value as criterion, which is calculated, not at the current level of fuel and lubrication materials prices, labour costs of flight attendants and flight crew, currency rates and airdrome fees stakes etc., but based on average predicted values over the total lifespan of an aircraft.

The obtained value of dynamic price value of a flight per one seat is of probabilistic nature and has a certain default level associated with predictive values of a certain number of indices. However, the accuracy of cost accounting in this case would be much greater, than when using static cost value of a flight hour, or the cost of either ton-kilometer or passenger-kilometer.

To carry out production capacities modernization at aviation industry enterprises one needs motivated buildup and approval of the investment policy. It includes the following basic stages:

1. The first stage — analysis of financial and economic activities of an enterprise, market outlet and competitiveness of manufactured products and assessment of production capacities condition.

At this stage, we determine the degree of moral and physical depreciation of basic production assets, and the line of activity selection to provide the necessary equipment including:

• Introduction of advanced techniques and technology;

• Mechanization and automation of production;

• Modernization and replacement of obsolete and physically worn-out equipment, and other directions.

2. The second stage focuses on the selection of the most significant projects, determination of the total volume of required investment resources. We recommend carrying out the analysis of the sources of investment resources with an allowance for the complexity of funds raising. The attraction of additional monetary resources with the use of credit and stock markets is proposed.

Based on the stated structure of investment capital for production capacities modernization, the determination of modernization of production capacities efficiency indicators, which the enterprise should approach, is realized.

Feasibility report of the project is included into economic mechanism of modernization of production capacities investment policy.

A special role in managing the implementation of the investment policy is assigned to the achievement of the required efficiency indicators of the implementation of production capacities modernization and the return on investment. For this purpose, it is expedient to carry out the justification of the need for investment resources considering economic indicators that reflect the costs of their involvement.

Calculations of the technology for justification of sources of investment resources application for production capacities modernization at the enterprises of the aviation industry efficiency include:

• Calculation of attraction of investment resources expenses, taking into account the time factor.

• Determination of the payback period of investment resources.

• Justification of the profitability of projects and calculation of additional profit as a result of production capacities modernization.

The execution of flights in the vicinity of an airport area in Russia differs from flights execution abroad. Before flight, Russian pilots set the airfield pressure on all barometric altimeters, the so-called mode of an airplane aviating according to the pressure over the airfield — QFE. Take off, After takeoff and circuit the altimeters are set to standard mean sea level pressure of 1013.2 hPa (760 mm. Hg) — QFH. This is the mode of airplane aviating according to pressure of the sea level. After taking a decision on the height of the flight execution of the flight at the set level begins. During preparation to landing after the plane descends to circle height, all barometric altimeters are set to the pressure over the airfield. After landing in QFE mode, the altimeters indicate the altitude of zero meters at the airfield.

Barometric altimeters do not change pressure indication all over the world world. While in flight at the level and at landing, barometric altimeters are set to standard pressure of 1013.2 hPa (760mm. Hg) and the whole flight is executed in the QNE mode at the height relative to mean sea level pressure. At the airfield, barometric altimeters indicate the remaining altitude relative to sea level. Pilots should know it, and take into account the location of every airfield relative to sea level.

The change of pressure in barometric altimeters results in errors and failures in interacting systems, such as system of automatic flight control, system of the early warning of dangerous closeness of earth of GPWS, system of warning of collision of airplanes of TCAS control the and signal permutation pressure due to height-indicators entered as barometric flight altitude. GPWS (EGPWS) provides radio altitude signaling. The change of pressure in barometric altimeters in the course of the flight is dangerous, and may result in a catastrophe at an attempt to go around in automatic mode. It happened in Kazan in 2013, in Perm in 2008 — at go- around of the airplane Boeing-737 in the automatic mode instead of circuit circling, when the set of height was translated to the autopilot. Eventually it resulted in a catastrophe.

For absolute promotion of safety, it is necessary to bring about the changes in the normative document of civil aviation of Russian Federation FAR-128 and oblige a crew to execute the flight on all stages only in the QNH mode.

The goal of the work is creation of the methodology of the solar-powered UAVs design parameters characterization.

The task of the solar-powered flight vehicle designing as well as any flight vehicle implementing aerodynamic flight principle comes to fulfillment of fundamental equations, that describe any projected plane as heavier than air flight vehicle with ability to fly in a certain altitude envelope and velocity range, make evolutions, take off and land at established class airfield.

Design problem of a solar-powered flying vehicle as well any other flying vehicle, realizing aerodynamic flight principle is reduced to design parameters characterization, which define any designed aircraft as a heavier-than-air flying vehicle able to fly in a predetermined flight altitude envelope and speed range, maneuver, take off and land at airfields of a certain class. The main of these relationships are as follows: mass balance equation; gravitation balance equation; power balance equation, as well as stability and balancing equation of an aircraft. In mathematical formulation of the problem, the design means meeting the above- mentioned requirements, considering the imposed constraints on either parameters, or characteristics. Solar-powered UAV design has some peculiarities due to the specific type of power plant.

The suggested methodology defines weight and geometry of solar-powered UAV depending on payload weight and power consumption. The methodology based on comparatively limited weight, obtained in consequence of gravitation and energy balance equation, and real weight, obtained from mass balance equation.

For more precise weight calculation, we create new weight models of aircraft elements. The aircraft parameters restrictions are included in the methodology due to atmosphere turbulence.

The methodology can be helpful as well in forecasting of solar-powered aircraft development if we take into account advanced technologies.

The paper considers an approach to developing pilot's control actions models during execution of typical flight modes. Implementation of such kind of models for the structures of pilot's backup systems will allow providing the possibility of parallel solution of task complex, aimed at pursuing an objective of a concrete flight mode.

The results of landing trajectories processing are reported. These results indicate that both aircraft trajectory parameters distribution and control parameters distribution during landing performed by various pilot's reveal statistic significant differences. This fact allows us to say that efficiency of upgrading of pilots actions support systems requires provision of possibility to adapt implemented models of pilots control actions, taking into account experience, qualification and peculiarities of control actions of a certain pilot.

We suggest an approach to forming individually adapted models of pilot's control actions based on neural network of a multilayer perceptron type. Such kind of model uses actual parameters of aircraft-pilot system status as input variables, including dynamic and control parameters. The output of neural network model considered in the paper is scalar indicator function, representing the convolution of parameters, characterizing the accuracy of aircraft touchdown to a runway. Parameters of the neural network model are determined as a result of neural network «training» by the data obtained during execution of previous flights. Thus, the suggested model allows predict the accuracy of bringing an aircraft to a runway based on current values characterizing aircraft-pilot system status.

The paper gives the results of pilot's control actions model building-up using the data obtained in the course of landing modes execution by two operators working at hardware-software MiG-AT aircraft simulator. Based on the obtained results we can draw the following conclusions. Adaptation of neural network model of a more experienced pilot is achieved only by updating weight factors of neurons with retention of its structure. With a primary pilot who does not demonstrate aircraft handling stable skill, adaptation of neural network model in the course of his professional activity is provided by changing both the neural network model structure and its parameters. Thus, the obtained results allow claim that individual manner of actions of a concrete pilot reveals either in the structure, or in parameters of neural network model, characterizing him.

We can get maximum effect from the use of unmanned aerial vehicles (UAV), if a certain infrastructure, including a group of UAVs, mobile control tools, system of commands and information receiving and transmission, as well as information- processing equipment is formed. The transition from the use of single UAVs to highly organized UAV complexes assumes the availability of advanced supporting and control complexes. Their creation is impossible without preliminary analysis and modeling of the processes of information interoperability of separate systems, included in such complexes.

The paper describes a conceptual model of complex monitoring system built by the authors, which uses UAV complex combined with both control systems and systems of information processing and delivery.

The object of modeling are information processes, occurring in supporting and control complexes: monitoring data acquisition, its transmission through the channel forming means and processing at various facilities distributed among the transmission path.

The goal of modeling:

— identification of successful strategies of complexes organization and functioning:

— operational feasibility checking of monitoring schedule, processing and connection under conditions of random interferences and hardware failure (or killing effects for military applications);

— the possibility of using the model as an instrument for monitoring current system conditions with the use of real data;

— the possibility of operational rescheduling caused by unexpected changing of system operating conditions.

Statistics obtained as a result of the imitation modeling enables evaluation of different approaches to organization and operation strategy of the system, and obtaining an starting point for more detailed, and hence more laborious optimization, reducing the search for optimal solution set.

The developed imitation model allows evaluate the effectiveness of monitoring system at the stage of design, and can be useful for monitoring (shooting) strategies selection.

The developed model was tested on a series of hypothetic examples distinct from each other in strategies of UAV use for fulfilling the task of surveying the territory with pin-point targets located on it. The modeling was carried out in the wide range of initial data and revealed operating capacity and non-contradiction of the developed imitation model.

Developers of software tools and algorithms for information interaction can be users of the imitation model. Besides them, the described model can be used for distribution of tasks among UAVs belonged to different launch teams.

In this work the standard software package on ANSYS CFX hydraulic gas dynamics is used for calculation of aerodynamic characteristics (coefficient of front resistance) of the rotating aircraft (A).

Currently, mathematical modeling methods are used for the determination of the aerodynamic characteristics for a wide range of aircraft. The main advantage of calculated methods is that once properly developed mathematical model and the program can be used repeatedly in different variants of the aircraft assembly scheme, which provides prompt problem resolution of determining aerodynamic characteristics.

This paper considers the simulation of the gas flow around the rotating smooth bodies based on accurate finite-difference method in a computing environment ANSYS CFX. The algorithms for solving the Navier- Stokes equations for the flow space, axisymmetric bodies by viscous gas on the basis of a complete system of equations that allows you to calculate the coefficient of drag of rotating aircraft for different Mach numbers and calculate the change in this ratio for the given angle of attack, which at the stage of setting the geometry of the computational domain is defined by the angular position of rotating aircraft relative to the oncoming flow.

Comparison of simulation results with field experiments performed, showed that the convergence in determining the drag coefficient is within the range of 5%.

Considered is the question of the influence of the rotational motion of the center of mass on the flight of uncontrolled aircraft. For that purpose, on the basis of the calculated data of drag coefficient of rotating aircraft with angles of attack other than zero method of least squares, allowed obtain the dependence of the ballistic coefficient of the angle of attack.

Thus, using the obtained dependence of the ballistic coefficient from the angle of attack it is possible to account the angular perturbations when calculating the parameters of the rotating trajectory aircraft.

Space debris (SD) creates more and more problems for the spacecraft operating in low earth orbits (LEO). At present experts worldwide come to the conclusion that there the need to «clean up» near-Earth space environment from space debris becomes urgent. However, now there is no adequate substantiation of active SD removal effectiveness impact on techmogenic pollution evolution. Thus, the efforts concerning the evaluation of active removal of large SD objects effectiveness impact on long-term technogenic pollution of LEO region are carried out within the framework of Inter-Agency Coordination Committee on Space Debris activities. The paper presents the results of the developed methodology as a part of these activities

The major element of the methodology is a statistical model of SD developed by A. I. Nazarenko. We suggest to describe the whole set of objects X as a set of objects described by three subsets: D — debris, ADR — the objects that can be subjected to the operation of the active removal, NEW — new launched objects with respect to which we consider technogenic pollution limitation measures. To solve the problems of mutual collisions and impacts of the collisions we shall describe the set X using a statistic SD model and multitudes specified in it. The subsets ADR and NEW will be described additionally using complete information on every object with a single-piece description. Prediction of orbital evolution of objects from the subset D will be carried out using methodologies of statistical model SD, the subsets ADR and NEW with prediction of the orbital parameters of each object. Accounting of active removal and SD limiting measures occurs while deleting objects from the subsets ADR and NEW, and as a consequence from the total set X.

Analysis of the results of implementation of the developed methodology showed that for technogenic pollution of LEO region reduction in the longer term at least 5 large objects a year should be removed. However, the impact of the effectiveness of the SD active removal on long-term evolution of SD would be estimated in the decades to come. The obtained results agree well with those of other researchers.

We identified the areas in which it is necessary to carry out the operation of the SD active removal in the first place:

1. h = 800 ÷ 850 km, i = 71,25°;

2. h = 650 ÷ 800 km, i = 98,75°;

3. h = 900 ÷ 950 km, i = 83,75°.

The paper considers the design of Thermal Control System (TCS) for TabletSat microsatellite. The satellite was launched into orbit (SSO, 606 km altitude, LTAN 10:30) from Yasny space-launch complex on June 19, 2014 for non-commercial technological experiments. The lead product developer of this satellite is JSC SPUTNIX. The satellite, just as the entire TabletSat platform, was designed according to Plug-and-Play principle. It required, apart from interfaces standardization, to use methods of fast thermal control system (TCS) development, which would allow provide proper thermal fields on-board for a wide range of payloads and spacecraft structures.

Due to narrow time and development cost constraints the decision to implement passive type of TCS was taken. For fast TCS development the authors recommend to fasten all side panels with standardized brass thermal bridges; connect all instruments and other thermal contact surfaces by heat-conductive adhesive; to install electric heaters with temperature control inside overcooling sensitive devices.

TabletSat-Aurora has the shape of a hexagonal prism, consisted of structural panels with onboard systems mounted on them, and deployable solar panels installed on the top panel plane. Panel optical enamel coating scheme is determined from computational modeling with modern CAE systems, such as Thermal Desktop, which the authors of the paper used.

Thermal modeling is carried out with allowance for heat currents from the Sun, the Earth and Sun radiation reflected form the Earth. It also considers the cyclic graph of on-board equipment heat emission. For each scheme of panel optical enamel coating the authors carried out time simulation for different spacecraft ADCS modes for 14 orbit passes, which compose full spacecraft operation cycle with payload energy-storage passes.

The best TCS structural scheme was sel ected according to the results of modeling. During the satellite flight operation data fr om thermal sensors mounted inside the onboard devices, batteries and solar panels have been collected. The experimental data agreed on the qualitative level with the results of modeling (the ranges of temperature variations obtained by modeling and in reality are the same). The authors determined also the results of future updating of fast TCS development standard for microsatellites, as well as recent trends of TM/TC system upgrade.

In accordance with the results of the satellite assembly, suggestions to structural elements update, such as search for optical non-crumpling enamel and replacement of heat-conductive adhesive by conductive gaskets to provide fast disassembling.

The solution of safety issues of gas turbine engines (GTE) for modern and advanced passenger aircrafts in emergency situations is a very important task of improving the reliability of aeronautical engineering. One of the possible and most severe accidents is a fan blade out of a running engine. In this case, the broken blade hits the engine hull at high velocity and may rupture it, which is impermissible. Numerical study of the blade and engine hull collision [1— 8, 11, 12] shows that the process of dynamic deformation takes place under complex stress conditions and characterized by high velocities and strain levels. Thus, we need reliable dynamic strength criterions to validate the engines impact resistance. When the blade hits the engine hull, its body deformation is local. Hence, to develop the dynamic strength criteria, we can examine a collision of rectangular plates, where local deformation is very similar to that of a real engine in terms of a stress state, deformation levels and rates. The aim of this work is to investigate numerically the impact deformation and develop computational and experimental criterions of impact resistance.

In this work, we examine the mutual collision of a flying rectangular titanium plate of a constant thickness with a quiescent aluminum target plate of a constant thickness. The target size was sel ected so as to eliminate boundary conditions impact on local deformations. The studies are carried out in velocity ranges V₀ ~ 160...239 m/s at different impact angles: φ = 90° (normal impact),φ = 60° and φ = 45°.

Experimental studies show that in the case of normal impact, at the impact velocity of V₀ = 160,8 m/s, the target is deformed, but not ruptured. At the impact velocity of V₀ = 195,0 м/с, the flying plate ruptures the target and gets stuck in it. For the impact angle of φ = = 60°, for two cases of similar impact velocities: V₀ = = 199,6 m/s and V₀ = 201,3 m/s, the target did not penetrate in the first case and completely penetrated in the second case. At the impact angle of φ = 45° in the range of velocities V₀ = 210,8...218,8 m/s the target did not penetrate.

The problem of numerical simulation of is formulated in Lagrangian representation, the equation of motion is based on the principle of virtual operation. The equations of state are recorded in the form of the flow theory with kinematic and isotropic hardening. The mutual collision of the plates is investigated numerically by solving a non-stationary contact problem with variable boundaries. Numerical simulation is based on the finite element method and explicit time integration scheme, implemented in LS-DYNA [9].

The results of numerical simulations and their comparison to the experimental data show that the duration of the active deformation processes equals t_a = 50...150 µs for all cases of loading V₀ = 160...239 m/s in the range of angles φ = 45°...90°. Maximum deformation rates develop at the impact side and reach ε = (5...7) 10³ с-1. During the mutual collision, in all cases the stress state close to triaxial compression П ~ —1,0 [10]) at the target impact side occurs, with biaxial extension П ~ +0,80 [10]) on the opposiste side.

The flying plate kinetic energy is not the criterion of target rupturing. With equal or similar kinetic energy of the hitting plate, the target destruction is largely determined by the conditions of the initial contact interaction («perimeter», «line», «point») that affecting the localization of the target deformations.

In all considered cases the target destruction starts fr om the impact side and is characterized by the local shear (cut) in the contact zone of the plate and the target. The deformation intensity stands for the measure of this shear and can be considered as the dynamic strength criterion. For the considered material, we determine the design and experimental criterion of dynamic strength as ε_i= 24...25%.

This paper deals with the problems related to the determination of the temperature state of contacting elements of aircraft engines, when contact thermal resistance occurs due to the imperfections of the mechanical coupling due to the discrete nature of the contact. It also considers physical aspects of the contact heat exchange and analyzes three components of thermal energy transfer.

The known analytical equations describing contacting in the air were analyzed in detail. The conclusion was drawn that the known dependences cannot be universal for practical problems, though the difference between the calculated values and the experimental data due to the uncertainty of numerous empirical coefficients and conditions of generalizing dependencies is essential. This statement is supported by the calculations with respect to the contact pair of stainless steel in a wide range of thermal and mechanical stress, which further allowed us to estimate additionally the contribution of different components of point contacts thermal resistance and environment.

Taking into consideration the revealed discrepancies and for the purpose of a more rigorous recordkeeping of temperature losses in the contact zone, the author suggests a fundamentally new generalized equation, which derivation is based on the classic analysis of the contact conductance components.

Describing successively the share of each value of the total thermal resistance in dimensionless form and justifying the influence of each of the values the author proposed the structural formula. The empirical coefficients of this formula are obtained by statistical processing of the well-known experiments. The results of summarizing are shown in the form of a graphic dependence of the contacting materials and their surface finishing.

According to the results of generalization a conclusion about a successful correlation of experimental and calculated data is made. The reasons of the marked differences are analyzed as well.

The results of the work with great certainty and precision allow us to determine the values of the thermal contact resistance between materials of special interest and to select their optimal values.

The task of development of a modern domestic unmanned aerial vehicle (UAV) with a cruising altitude of 18-20 km and characteristics equal to the best foreign prototypes is very important. The requirements relating to such type of UAVs are still not determined not only to the aircraft complex as a whole, but also to its individual systems, and power plant (PP) in particular. Thus, a wide range of urgent issues arise in this area requiring research and working out, including the design of engines of new generation for PP of such kind of UAV.

The authors suggest in this respect a methodology of forming a PP preliminary technical appearance for a high-altitude UAV. It will allow selection and substantiation of rational parameters of the operating procedure, controlling program, scheme of gas turbine engine (GTE) of straight reaction, as well as characteristics of the PP inlet and outlet devices, guaranteeing the achievement of the extreme (best) values of the UAV performance factors.

The key core of this methodology is the integrated mathematical model (IMM) of the technical system Unmanned aircraft — power plant, developed by the authors. This IMM is intended for carrying out various exploratory parametric and optimization studies on the formation of technical characteristics of various schemes of turbojet engines, as well as evaluate PP internal parameters effectiveness and UAV performance criteria, such as, its performance characteristics.

The authors carried out the assessment of adequacy and accuracy of the developed IMM of the UAV-PP system by the example, using a number of existing engines and aircrafts. As a result it was concluded that this IMM is accurate enough for conducting various engineering studies concerning the formation of a preliminary technical aspect of the aircraft GTEs of straight reaction. Using the IMM we conducted a series of parametric studies of PP with bypass turbojet engine parameters impact on its characteristics and UAV characterixtics

This work formulates and solves optimization problem, consisted in the best PP technical aspect selection by the performance criteria of the UAV — maximum range and duration of the flight. As a result, we obtained Pareto set of optimal solutions, from which a compromise variant of UAV-PP system, ensuring the improvement of aircraft performance characteristics as compared to a certain basic (non-optimal) variant by flight range of 10.4%, and duration of 9.3 % was obtained.

The practical value of this work consist of the following. Its results can be used by scientific and engineering organizations, engaged in the design of perspective UAVs and their PPs, as well as by the air force and in industry for validation of the requirements to new patterns of aeronautical engineering, as well as by aircraft engineering institutes to improve the educational process.

This paper considers the role and place of mathematical model of a turboprop engine power plant mass generation in the problem of gas turbine engine operating procedure optimization at the stage of initial design. The mass of a power plant with turboprop is one of the major components, actually defining parameters matching of an aircraft and its engine, in the course of forming a takeoff weight of an aircraft, which indirectly defines the costs of material resources to create an entire aircraft designed system. Moreover, in the performance of parametric studies the project weight of a power plant should be estimate by turboprop parameters. However, such dependencies hitherto are underresearched. Thus, the evaluation of the engine weight dependencies on its operating procedure parameters is carried out based on either integrated data on generalized statistics on completed designs, or on weight parametric models, as far as more precise information is not available at this stage. We compared a number of state-of-the-art approaches to evaluation of the mass of a power plant at the initial design stage. We proposed also several variants of power plant mass models, differing from each other by the number of parameters required for calculation at the initial design stage and accuracy of the models themselves. A comparative analysis of mathematical models reveals their advantages and disadvantages for different stages of the design.

The paper presents designing computation and defines specific energy-mass characteristics of a new type of power plants (PP) for off-line power supply of an astronaut in space, based on oxygen-aluminum O2/Al chemical current source (CCS).

Such PP with 100 W rated power of 27 ± 3 VDC and short-term current change from 0.45A to 35A should provide necessary energy storage for active operation during 180 hours (30 six-hour cycles) with long-lasting pauses, as well as storage capability within a year before setting to operation.

The authors developed structural diagrams for 1, 5 and 15 six-hour cycles. The simplest PP structure and operation mode demonstrates the PP meant for one- cycle. It almost fully autonomous in relation to on-board systems and highly reliable. However, to support the entire program we will need to install 30 PPs of such kind on-board a spacecraft. After mission completion such kind of PP either recycled, or returned to Earth for recharging.

PP designed for five operation cycles with prolonged pauses has more complicated structural diagram. After operation cycle completion, this PP is connected to conservation system, and should be refueled with Oxygen before the next start-up. However, only six of such PPs are necessary on-board the spacecraft, instead of 30 pieces of the previous kind. The structural diagram of O₂/A1 PP for 15 operation cycles is most complicated, since it includes addition electrolyte adjusting systems, as well as its solid resultant (aluminum hydroxide Al(OH)₃) purification, and consumed water supplying.

Comparative mass calculations were performed for all PP variants. We choose the minimum PP mass as an optimization criterion, and the value of current density in O₂/A1 elements as a variable parameter.

Calculations were performed on experimental data obtained with the O₂/Al CCS laboratory sample with aluminum anode, made of Aluminum-Indium alloy (A995 + 0.6 wt.% In In0), in alkaline electrolyte based on 4M NaOH with addition of 0,06M sodium stannate Na₂SnO₃, that inhibits aluminum corrosion.

As the calculations showed, the installation designed for one-cycle operation has the minimum weight. However, the PP designed for 15-cycle operation exhibits minimum total weight and best specific characteristics.

It was shown for the first time that hydronic CCS with aluminum anode may also be used in space as the hydrogen generator for the hydrogen-oxygen (O₂/H₂) fuel cell (FC) in the combined PP (CPP). It solves effectively and safely the problem of prolonged storage of hydrogen for autonomous PPs based on O₂/H₂ FC. Energy characteristics of such CPP exceed those of PP consisted of only of O₂/H₂FC. When put into practice such CPP will be able to satisfy the requirements on long pauses between runs and may be considered as an autonomous energy supply for an astronaut.

The author first proposed to consider the process of thermal degradation of polymers as a kind of process of physical-chemical relaxation. The process has two components: chemical, due to the rupture of chemical bonds and physical, associated with the movement of the structural units of the polymer. In this case, the so-called apparent or effective, kinetic characteristics of the process — the pre-exponential factor, activation energy and reaction order — purchase a specific physical meaning. Easy to overcome the energy barrier of a chemical reaction is preceded by the movement of the respective nanostructured polymer units, that is ongoing in parallel with other processes, with nature and relaxation which also requires a certain energy.

Calculation of kinetic parameters is important not only for the formal description of the process, but also to obtain data on the structure of the polymer. For this purpose it was proposed to consider the process of thermal degradation of the PM as kinetic plane described by the three parameters. Defining segments, that the kinetic plane trims on the coordinate axes X, Y, Z, and the angles formed, it is possible to calculate the kinetic parameters of thermal degradation process of the polymer.

The author introduced the concept of the relaxation time of thermal degradation of the polymer τ_р. Asked to count it as a time of transition of the polymer fr om the initial equilibrium state to equilibrium state corresponding to the high-temperature heating conditions: ( T_f—T_i)/b wh ere T_i and T_f is the temperature of the beginning and end of the thermal degradation process, b is the heating rate, K/s.

This approach allows us reliably interpret experimental results and to predict the thermal behavior of polymeric materials at high speeds heating. Specialists receive information, allowing the use of nanotechnology for targeted development of the nanostructure of polymer materials heatproof destination. On the other hand, the creation of nanostructured polymers and polymer composites is achieved processes of thermodestruction management. Of special interest, relaxational nature of thermal degradation of polymers should cause researchers stress-strain states of polymeric composite materials, working in conditions of high temperature heating and used to create structural elements of modern aerospace.

The article describes methods of winding ratio determining in AC electric motors with tooth stator coils and permanent magnets rotor, and presents the values for several design styles of such motors.

On the example of the three-phase two-piece electric motor with radial flux, 18 teeth of the stator and 10 rotor poles, an analytical expression of the winding ratio was obtained (from the slot e. m. f. first harmonics vector diagram and the instantaneous rotor position). Additionally, alternative analytical method for determining the winding ratio is shown (based on the known expressions for the motors with distributed two- layer stator winding with adaptation to the tooth winding motors).

As the result of this calculation it we concluded that the value of the winding factor increases with the number of coils in the coil phase group.

The paper presents a comparative analysis of the electric motors of distributed and tooth stator windings design, highlights a number of positive features of tooth windings motors. We made a hypothesis on eddy currents origin in rotor elements due to the presence of pulsating electromagnetic stator flux. Distribution nature and eddy current losses value that occur in the rotor and its elements (for three different speeds) are shown with the motor model in Ansys Maxwell v.16. package finite element analysis.

As one of electric motor efficiency factors, the value of winding coefficient (obtained for a particular number combination of the rotor pole pairs and stator teeth) allowed choose the most efficient motor in electromechanical actuator design stage. Knowledge of eddy current value and distribution nature in the rotor design elements allowed take measures, leading to these losses reducion.

The article is prepared in accordance with R&D work High-tech production creation of precision high- speed enforced electromechanical actuators of the new generation at ITMO Univercity, with financial support from Russian Federation Ministry of Education and Science.

Modeling of the electromagnetic environment is a promising direction of designing the systems on board the spacecraft.

Modeling electromagnetic environment allows you to define the levels of radiated electromagnetic interference in different parts of the spacecraft structure and draw conclusions on the need to improve on-board equipment and changing ways of laying cables.

The paper presents a method of electromagnetic environment modeling of systems on board the spacecraft. The method consists in replacing the real on- board devices, emitting electromagnetic interference (EMI), by equivalent sources. The advantage of the method is that it allows recreate electromagnetic environment created by on-board devices and a small number of experimental measurements. Studies of real samples of on-board devices can significantly improve modeling accuracy of electromagnetic environment due to the individual features of electromagnetic interference in electrical circuits of each of the devices. This method allows model electromagnetic environment according to geometric dimensions and arrangement of devices anywhere in the spacecraft.

The paper gives examples of modeling of on-board instruments on the basis of the proposed method. As imitators corps on-board devices used metal construction of a cubic structure with attached generators. Initially, the electric field strength is measured using an antenna, then calculates the values of the field theoretical method. The modeling results are presented in the form of the spatial distribution of the electric field.

New circuit approach to aircraft on-board and airfield electro-pulse anti-ice systems with high reliability, electric power efficiency, assembly processability as well as electromagnetic compatibility is examined.

Anti-ice systems represent a key facility for flight security and aircraft installations normal operation. They play an important part providing successful operation of following. They provide successful operation of the following on-board & ground-based equipment: air-field radar, light signal, high voltage, wind-driven, antenna feeder, gas-piping and etc. (especially for facilities that are operated out-door under winter conditions or in the north regions).

Electro-pulse anti-ice system (EP AIS) represents a certain variety of mechanical anti-ice setup. It provides repetitive pulse impacts in a protected cover. At present the electro-pulse anti-ice systems based on remote contactless impact on the cover is used in aeronautics. They induce pulse eddy currents by means of dedicated EM inductors. These eddy currents induced in certain cover surface areas interact with magnetic field of the latter and produce the pulse vibrations of the cover, and, split off of an ice layer. The debris carried away by the air stream.

The main EP AIS advantage is its low power consumption — ten and even hundred times less than other systems energy consumption.

The general shortcomings of such systems are as follows: low system assembly processability due to bulk inductors for various parts of the cover (their area of action is limited by surface junctions and device reinforcing elements), residual ice formation under relatively large exposed area; the necessity of pulse power boost in case of higher rigidity of construction; structural complexity and its low reliability for composites.

There is a reasonable way to reject the induction vortex interaction between vibrator inductor coil and the cover: the usage of the pair of cores — fixid and movable one. The latter is fixed on the spring and is supplied with a percussion lug for knocking the cover (including materials with low electro-conductivity -composite via metallic substitutes).

The power stage of the secondary pulse power source for anti-ice vibrators with primary supply from low- voltage DC mains network (27 V) is presented in this article.

The circuit represents a reversible (bidirectional) step-down / step-up pulse converter with reactor integrator and a damper-snabber non-dissipating circuit for modulator switch «soft» commutation.

The power taken from the power source is relatively small due to energy batch recuperation from reactor integrator back to power source and to input C- integrator as well as due to commutation loss reduction by soft key switching. In addition the switch commutatin overstress and inversely-diode overcurrent are eliminated. Moreover, the HF commutation noise emission is also reduced.

The new circuit solution examined in this article is intended for on-board and airfield electro-pulse anti- ice system design. It provides the equipment functional capabilities enhancement due to bidirectional energy conversion, particularly its recuperation, as well as efficiency and reliability enhancement, EM compatibility improvement thanks to switching overstress and inversely-diode overcurrent elimination as well as switching heat loss

The presented circuit solution is protected by Russian Federation priority. The research presents interest for a wide circle of specialists in the field of aircraft electrical equipment design as well as anti-ice equipment for ships and stationary installations, used for instance in airfield radar and lighting equipment in the north.

The paper is devoted to the issues of implementation of the concept of accelerated import substitution by the example of aircraft engine building. Despite the fact that Russia has created a powerful industrial and scientific potential, it does not provide the proper competitiveness of aircraft industry. The situation is near critical since the civil aviation market is lost. For the most part, indexes of effectiveness of home-produced aircraft engines should be adjusted to international standards. It is proved that high efficiency of import substitution of this industry sector can be achieved only using complex approach, based on real and rationally organized integration of science and industry, when manufacturer works in interactive mode with scientific departments and organizations in the course of design of new samples. It also requires upgrade of integrated vertical structure on the principle of cyclical turnaround and horizontal links; provision of in-home cooperation effectiveness; innovation projects and programs, methods and technologies; comprehensive planning and marketing management; high quality and ecological properties assurance; building up unified virtual environment of design and engineering analysis; support of pilot and mass production, after-sale services, inter-industry integration.

The goal of the state policy of import substitution in this situation means creation of favorable conditions for the coordinated development of the multi-project environment, which will reduce or neutralize the risk of uncertainty growth and take into account various effects. Russia is the one of only five countries that can independently develop and produce aircraft engines. Since 2008 Russian airlines have acquired nearly 600 foreign aircraft, and only 59 produced in Russia, while the share of modern Russian aircraft account for about 7% of the total fleet of Russia, and their contribution to total passenger traffic volume is only 4%. Intolerable situation in the park of medium-haul aircraft, where the share of foreign aircraft is about 80%, i. e. it is about the Reconquista own market. In the production of short- haul passenger aircraft SSJ-100 the share of imported suppliers is about 60%. Now you need to bet on the innovation and quality of the engines. Thus, the main directions of improving the competitiveness of the industry include:

• drastic improvement of performance, reduction of noise and emission of harmful substances;

• cost reduction in the production and operation of aircraft engines;

• development of critical technologies with the advanced breakthrough technical solutions and technologies.

In order to improve flight performance the most problematic issue for domestic aircraft engines is achieving optimal values of the following indicators:

• the level of reliability of the power plant;

• fuel efficiency;

• the cost of maintenance of aircraft engines;

• maintainability;

• compliance with future environmental requirements.

The main directions of implementation of programs aimed at costs reduction are as follows: minimization of losses; outsourcing of services; productivity increase:

• upgraded vertical integration of the productive cycle and strengthening horizontal linkages;

• ensuring effective inter-organizational interaction.

The substitution in aircraft engine, as a strategically important field, can be effectively achieved only as a result of sound economic policies, creation of the next generation of products based on complex approach, which provides a logical linking of the trajectories of development for individual enterprises producing aircraft engines and other aviation equipment and related industries and activities.

The fast import substitution strategy should be considered as a set of concurrently running projects focused on competitive domestic and foreign markets, for which it is necessary to create favorable conditions for the coordinated development of the multi-project environment.

Accelerated substitution is a public task, which is possible only under condition of use of the integrated approach, where special attention should be paid to financial support in order to establish scientific and technological basis for critical technologies, components and systems for advanced engines and increase their advanced pilot producing during the stage of research.

Uncertain demand for output products forces a manufacturer to make planned decisions under conditions of uncertainty and risk.

An efficient tool for enhancing the quality of planned solutions is implementation of the games theory methods. The article analyses the task of searching for an optimal decision in a situation, where the volume of government-guaranteed orders is a source of uncertainty and risk.

A search for optimal decisions comes down to the game with nature, in which one player — companys management — acts consciously and the second player — nature (state customer) — makes decisions regardless of the companys management and is guided only by changing of external economic and political conditions.

The article considers also methodological issues related to making a decision on optimization of a Government order performance plan by the example of a defense industry enterprise.

The article describes the key elements of payoff matrix related to the game with nature. The companys profit maximization is a criterion of optimal planning decision.

The maximax criteria advanced by Wald, Hurwitz and Savage under conditions of uncertainty is applied.

Applying an expert probability estimate of the state of nature (a volume of government orders) makes it possible to calculate the risk criteria by Bayes, Laplace, Hodges-Lehmann. The opinion consistency within a group of experts is estimated by Kendalls and Babington Smiths coefficient of concordance.

A comprehensive comparison of total calculated criteria allowed developing an optimal strategy for preparation of government order performance plan ensuring the companys maximum average profit provided that it is fulfilled within the established time limits.

The goal of development and implementation of the assessment of in-house staff training effectiveness of a production enterprise is the increase of the return on the stuff of an organization education investment by assessing and identifying resources to improve the socio-economic benefits of training, retraining and skills development.

Design/methodology/approach

To create this methodology the author developed a classification of approaches to efficiency upgrading of training, retraining and professional development of the stuff: investment, psychological and competency, institutional. As far as the assessment of the result should contain economic and social aspects, which results are internally interdependent, economic outcome is socially significant, and achievement of economic performances creates objective possibilities for more complete satisfaction of material and moral needs.

Conclusions

Implementation of the methodology for assessing the effectiveness of training at production enterprise will reduce not only time, but also the financial resources of the organization. In general, professional development plays an important role, since in the course of the learning process the worker acquires a new theoretical knowledge required for the implementation of labor activity. It increases the ability of the personnel to adapt to changing economic and working conditions as well.

Practical recommendations

The methodology of assessment of the in-house stuff training efficiency at a machine-building enterprise is developed, differentiated depending on the purpose of assessment. It consists of:

1) effectiveness assessment by stuff levels — a separate employee, department, organization;

2) effectiveness assessment by stuff categories: workers, specialists, executives;

3) determination of economic, social effectiveness, the degree of meeting a set goal objective.

Originality / value

The methodological approaches for improvement of in-house training and evaluation proposed by the author can improve its effectiveness. Investment, psychological, institutional and competence approaches contribute to the optimal planning and content of the educational process, determining the effectiveness of training and identifying the main factors affecting the in-house training in general.

Subject area of the case

Business economics of aerospace industry enterprises.

Method

The studies carried out on the analysis of the current phase of the space industry reform bear generalized popular- science character. The article analyzes the preliminary results of the corporate activities of rocket and space industry enterprises in 2013 and provides suggestions for optimizing and building effective corporate governance, considering among them the possibilities of expenditures minimization in its practical incorporation.

Results

Information obtained in the course of preliminary work as well as carried out analysis of statistical data indicate that ongoing works on reforming the space-rocket industry with regard to structural transformations need some adjustment for the purpose of forming mechanisms of personified responsibility for taken decisions and final actual result.

Practical implications

One can use the results of the study to build an effective system of corporate governance in the existing integrated space-rocket industry structures, as well as to search for new sources of domestic investment space-rocket companies included in these associations.

Conclusions

Currently, it is possible and necessary to use the existing potential of the Federal Law of 26.12.1995 № 208-FZ On Joint-Stock Companies for the optimization of internal corporate communications in order to find sources of domestic investment companies of aerospace industry.

The article discusses the methodology of restructuring the aircraft manufacturing enterprises based on their division into business units and allocation of industrial park at vacant production areas.

The authors identified three groups of methods of manufacturing assets of aviation enterprises restructuring, and suggested the ways of carrying out the restructuring of aircraft manufacturing enterprises production assets according to the selected groups.

The article presents the performed analysis of the reasons for the unprofitability of the enterprises of aviation industry and determines the groups of factors, which cause it: design-engineering, economic, financial, in-house governance. To eliminate the causes of unprofitability we suggest a package of measures concerning utilization of free production facilities. The basic measure is the restructuring of production assets based on the integration of the plant and industrial park.

For restructuring of production assets of enterprises of the aviation industry we considered normative and legislative acts making provision for such restructuring. We considered the following concepts:, technological park, technological park in the sphere of high technologies, industrial park, business park, technopolis, cluster (technology area), and revealed the differences between these structures.

Examples of creation of technological park at the aircraft-building enterprises are given.

Such concepts as anchor and invited residents are introduced.

We suggest a technological park management structure based on financial accounting centers (FAC), and allocation of functions between FAC and technological park executive heads. We outlined as well the functions of the treasury (accounting and finance center), and showed the ways for increasing the profitability of a profit-center. We assigned to such methods: cost saving, functional and cost analysis of the products to the effect of cost reduction and increase of income and bonus.

The article suggests allocation of functions between FAC and avionic technological park executive heads, and highlights these functions.

The article considers the main requirements to a research and development cost management system at the aircraft engine-building companies. These requirements are claimed both in terms of research and development features as an activity category, and relative to the aircraft-engine-building branch peculiarities stipulated by specifics of life cycle of gas-turbine engines. We determined the main limitations imposed on the research and development cost management system with a view to the necessity to carry out the existing norms of active legislation, implemented to ensure State defense order. Based on an information available in free access the paper gives the examples of practical suppositions of the necessity of research and development cost management at aircraft engine-building enterprise JSC NPO Saturn.

The paper considers the classification of cost management methods according to organizational and administrative level of applied approaches. Based on the considered requirements and limitations we suggest recommendations on development, organization and selection of trends of development of research and development cost management system at an aircraft engine-building company.

The research and development cost management system should provide sustained management in conditions of:

• uncertainty and risk;

• existence of a set of diverse projects.

Cost research and development management at the aircraft-engine-building companies should be carried out by the integrated subsystem of business management providing rational distribution of resources between existing projects according to the purposes of the organization, and representing generally system of norms and the rules supported by mechanism of their execution.

One of the variants of classification of cost management methods is the group of cost management methods depending on organizational and administrative level on which administering is exercised. By following this classification, it is possible to split cost management methods into three main groups: operational methods, managerial methods and strategic methods. Examples of popular cost management methods are target costing and earned value analysis. These methods are widely used in the sphere of management of research and development and in project management.

It is worth noting that only reasonable application of various methods of cost management at operational, managerial and strategic levels will allow minimize disadvantages of applied approaches, and achieve synergetic effect from a combination of their advantages. Methods of target costing and earned value analysis considered in the present article are widespread methods of cost management at the enterprises and have considerable theoretical and practical base of use. Nevertheless, their full integration into an enterprise management system is still involves considerable difficulties and related to resistance, both from managers, and from the staff.

Increase of efficiency of use of methods of target costing and earned value analysis is obviously possible by organic integration of these methods into the intra- corporate institutes providing comprehensive support of adoption of administrative decisions at all administrative levels.

The article presents the executive summary of risk controlling of the continuity of productive processes and mechanism as applied to the enterprises of the aircraft engine industry. Creation of world-class corporations in the key segments of aviation industry requires, among other things, design of aircraft engines that meet international standards and allows a successful entrance to the global market of domestic corporate formations in the aircraft industry, as major players for long-term perspective. The necessary condition for stable functioning and development of aircraft engine-building enterprises is their ability to detect and prevent in time negative consequences that can jeopardize the very existence of the enterprise. In aircraft engine-building industry, as in many other branches of Russias military industrial complex, the necessity for radical restructuring of the enterprise management system is urgent, and under conditions of uncertainty in particular. We need to introduce new tools of enterprise management by which the continuity of production is provided. One of the tools responsible for ensuring the stability and survival of an enterprise is the continuity of the production processes risk controlling.

Analysis of the research in the field of management, productive processes continuity at the enterprises of machine-building industry both in Russia and abroad has allowed development of conceptual basics of the productive processes continuity risk controlling for enterprises of the aircraft engine-building. The article presents the structured analysis of the existing standards for continuity of production activities allowing develop the principles of the productive processes continuity risk controlling.

We classified and substantiateв the data acquired in the course of the study. It allowed developing scientific methodological issues of developing productive processes continuity risk controlling system for aircraft engine- building enterprises. Based on productive methodology of activity continuity control we developed and tested the mechanism of production continuity risk control at JSC Kuznetsov.

Risk controlling is not yet widely used at Russian machine-building enterprises. The scientific research in this field is also staying at the state of growth. Moreover, it does not exhibit systematic approach to the development of productive activities continuity risk controlling. Thus, it is advisable to bring about risk controlling to manufacturing control, using foreign and domestic experience of advanced corporations.

At present, there is an urgent need to re-equip and rearm of the Armed Forces of Russian Federation (Russian Armed Forces). Scheduling and control of the state defense order (SDO) are important and of government priorities. Formation of requests in the state defense order is a complex and multi-task issue. Its results affect the great number of people, many areas of military service and combat readiness of the Army as a whole. As practice shows, not sufficient time is allowed for the formation of the proposals for the SDO. Moreover, due to lack of personnel and time, SDO is often planned based on the current situation, focusing on the SDOs of previous years, either by leaving the last years numbers, or increasing them by a certain factor. Without automation of the process of the needs of the troops estimating for a specific planning period, the formation of the SDO is superficial and does not reflect the true state of affairs that is unacceptable in such animportant task as the defense of the country.

To automate the process of developing the plan, the state defense order for a certain planning period was scheduled using the following technique.

A method of short-term forecasting was sel ected and Excel was used as a simulation tool.

The source data includes information on certain types of troops supply with weapons and military equipment in a given year. It is based on the reports fr om the troops.

According to the results of calculations it is possible to build charts and diagrams, representing in a convenient visual form various development strategies, state technical means at a particular level of funding, which allows you to adjust the financing of the state defense order and specific technical systems.

With a minimum set of hardware and software it allows you predict the development, procurement and repair of technical equipment and weapons for the desired period of implementation of the state defense order with quick and accessible representation of the obtained results to the interested military authorities.

The article considers wing design parameters, having the greatest impact on its inductive drag. It becomes up- to-date due to the fact, that with low speed flying vehicles, and paragliders among them, inductive component of a sing drag appears to be several times greater than its airfoil part component.

From Expressions obtained show that with flying vehicle specified mass and its vertical overload the value of inductive drag of its wing will be inversely proportional to the square of its wingspan and square of its speed. Thus, with equal wingspan and speed, either narrow wing with high aspect ratio and high-lift airfoil, or wide wing with low aspect ratio and low-lift airfoil having equal lifting force will possess the same inductive drag.

This implies also the futility of replacement of wide wing by several wings with a number of narrow chordes. Even without considering polyplane wings negative interaction with total area equal to paraglider wing area and wingspan, polyplane wing inductive drag will not be less than that for a monoplane. It is shown for gliding systems, how we can determine required C_ya by C_xa and gliding speed, or find gliding speed by aerodynamic coefficients in terms of gliding system characteristics. In both cases we can calculate glide angle and corresponding to it system aerodynamic efficiency in terms of wing parameters; and then locate it relative to shroud lines in such a way that provide required wing incidence angle during gliding.

Any two airfoils having the same Kmax, create the same drag at a given lifting force, but the airfoil of the same quality with greater lifting force will require smaller wing chord and, accordingly, greater aspect ratio and minor stiffness with one and the same wing inductive drag. Besides, taking into account that extra lifting force value under lift devices deviation is proportional to wing area, we are interested at maximally wide chord. Thus, airfoils, providing minimum C_ya with the same aerodynamic efficiency, are preferable.

It is worth mentioning, that the demand for atlas of special airfoils for soft wings in the frame of aerodynamic quality, load-bearing capacity in deformed state, as well as safety, still remains impelling and unsatisfied.

Parameterization algorithms for main structural elements of helicopter BMR, i. e. bearingless main rotor, (including rigid rotors) is a multiphase task.

The first stage involves the static calculations and construction of blade resonance diagram.

The calculations of alternative stresses is the content of the second phase of the parameterization.

The calculations of the local strength and stiffness is the final phase of calculations based on finite-elements model.

This article is devoted to the first stage of BMR parameters selection algorithm and BMR durability analysis. Its main results are also used at subsequent stages of calculations and design.

The first stage contains calculations of rotor hub stiffness matrix, blade free vibration modes and frequencies, construction of blade resonance diagram, blade deformation calculations and constant stresses caused by centrifugal force impact during the flight, and blade droop caused by gravity when the helicopter is on the ground.

To provide static strength, stability (self-oscillations of flutter type) and dynamic strength of bearingless main rotors its essential to use more complicated calculation models and methods than those used for conventional hinged rotors, due to the complexity of the boundary conditions, used to solve differential equations of natural and forced oscillations of a blade. The paper proposes the solution of this problem by introducing hub stiffness matrix to analytic model, and offers the methodology of its elements calculation.

The article describes the design variant of light helicopter bearingless main rotor and presents the results of of BMR rotorhead main parameters selection, which can be generalized for a wide range of designs.

Sometimes interaction of elastic transverse vibrations of the body with an incoming air flow is referred to as aeroelastic vibrations. The frequency of these vibrations typically is close to the body proper vibrations. The vibrations induce unwanted transverse inertial loads on the body, particularly when they reach high amplitudes.

In the case when the missile is represented as a perfectly solid bit or when it has a rigid body, only the small perturbations of the motion parameters are considered. It allows to obtain with regard to perturbations a linear system of differential equations with coefficients that depend on parameters of the unperturbed motion.

Small thrust vectors induced by transverse vibrations of the body with accurate within the second order of smallness result in no variations of the missile acceleration in longitudinal direction. This acceleration remains the same one as for the nonperturbed motion both in value and in direction. This is the reason why under the small perturbations the transverse motion can be considered regardless of the longitudinal motion. Within a short period of time the guided missile trajectory can be considered as a plain curve, which is little different from a straight line. Therefore, let us consider the nonperturbed motion of the missile within a short period of time as the straight-line motion.

The issues of influence of flexible oscillations on flying ballistic and strength characteristics are arising during the process of designing missiles with bigger lengthening. Flexible oscillations occur from the engine tractive force and aerodynamical forces.

In this article the missile is substituted for inhomogeneous pivot for the purpose of studying the flexible oscillations.

Consideration of missile as inhomogeneous pivot significantly simplifies the study of angle characteristics determination.

The differential equations of the missile motion with regard for flexible oscillation have been derived.

The differential equations of the missile motion include torque and forces of the generalized coordinates of transverse vibrations are first sounded.

In the presented work such characteristics as a tractive, ballistic, weight, aerodynamical and inertial one are substituted for average ones during missile operation.

First tone flexible oscillations are taken into account.

Analytical dependencies of slip angle change, speed vector change, missile longitudinal axis angle with regard for the flexible oscillations have been obtained.

The article is devoted to simulation of a spacecraft thermal control system with the use of new simulation paradigm. The paradigm is based on three-stage decomposition method, which has been proposed by the authors. The thermal control system under consideration follows the same principles as the on-board systems of real spacecraft like Soyuz TMA-M and Progress-M.

At the first stage the hierarchical decomposition is made according to the spacecraft dividing circuit. As a result, the system is divided into parts, which models are named components. Components are nested into each other and establish a hierarchy similar to the dividing circuit.

At the second stage the interface decomposition is made when interface entities are picked out in every component prototype. Models of such entities are named simple models (SM).

At the third stage the interactional decomposition is made when each SM prototype is considered as a member of various physical interactions. Each SM is composed of set of fragments that represent the SM prototype involvement in interaction of specific physical domain. Each fragment is an object that contains the SM prototype parameters related to the physical domain. Additionally, the number of subroutines is assigned to each fragment that calculates relations between the parameters.

While modeling the various links between parameters of different fragments of the same SM are taken into account. That links are named cross-fragment links and implemented with particular subroutines each of which corresponds to the unique fragments combination.

Simulation model structure is defined with particular objects bonds. Each bond joins two SMs. Bonds declare copying of parameters between fragments of linked SMs.

In this paper the creation of universal standard elements library is considered. The library includes, firstly, the fragment classes definition, which are used for simulation of fluid flow through spacecraft subsystem pipelines, secondly, subroutines for simulation of relationship between heat and hydrodynamic domains, thirdly, bond definition, and, finally, the description of component classes used for spacecraft thermal control subsystem simulation.

The main feature of the library is a wide use of cross- fragment links that makes the flexibility of designed fragments possible. For instance, any pipeline for gas flow including heat transfer as well as for liquid isothermal flow is simulated with the use of the same fragments that define pipeline geometric properties. The flow of particular substances is simulated using, firstly, special fragments that define substance properties and, secondly, cross-fragment links between fragments of hydrodynamic, thermal and substantial physical domains.

By means of this library the simulation model of a spacecraft thermal control system was designed. The results of the typical processes of simulation have been obtained; the conclusion about the correspondence of the calculations results to the real spaceship telemetry data has been made.

The library can be used for simulation of real spacecraft thermal control system with a similar design and operating principle. The library may also be expanded for simulation of various flow processes of different substances of various states, what allows to simulate another spacecraft onboard subsystems like life support system.

The quality of aircraft engines (specific consumption, gas-dynamic stability, reliability, service life) is embedded in the course of the design process, fixed outat the stage of finalization, providedin the course of production, and implemented during operation. Account for actual geometry of the partsduring design process, and forming the best for current manufacturing environment conditions geometry of the products during manufacturing and assembling is one of the reserves allowing aircraft engine quality increase. Realization of the above mentioned reserve becomes possible with implementation of intellectual system of aircraft engines quality provision. The article presents the technique for the intellectual system of aircraft engines quality provision forming. It shows the structure of the intellectual system of aircraft engines quality provision. Mathematical model of the system with regard to analysis and clustering of complex shapes and surfaces is built. The analysis allows calculating parameters of a part location deviation from nominal geometry, and calculation deviation of parts shapes. Analysis tools are as follows: an iterative algorithm of nearest points, piecewise splines approximation of surface points, Fourier transform. We used an algorithm of k-means and self-organizing Kohonens neural networks as clustering methods. For adequate profile parameters clastering (shape and position) these parameters were normalized to the range of [0,1]. Complex shape geometry intellectual analysis model was implemented in MATLAB^®. Within the framework of the developed system using obtained models, clustering of series of GTE backet compressor blades, obtained as a result of measuring with CMMs DEA Global Performance 07.10.07. Implementation of neural networks seems to be more promising for the tasks of clustering solving compared to static methods, due to the possibility of learning and clustering of new objects on the basis of accumulated know-how. For further model realization in manufacturing building of new program modules is necessary

In terms of the guidance system of the guided aircraft air-to-air missile methodic prerequisites of the assessment of the technical state of one class of complex technical systems were suggested. The assessment of the technical state of the guidance system of the guided aircraft missile commentary was carried out using two methods. The first one was the parameter control method in the universe of linear nonqualified tolerances. The second one was the verifying method on figure of merit in the functional-hardware potential field. The presentation of returns of control was deduced using both accounted methods in the space of parameters specified a combat potential of the guided aircraft missile.

The method of the assessment of the technical state of complex technical systems and its instantiation inside of the intellectual technical diagnosis system were in [1- 4] established. In this article the development of the above mentioned approach for the assessment of the technical state of the guidance system of the guided aircraft air-to-air missile was reproduced.

In [5] the hybrid control method was offered. Within the framework of this method the current technique of parametric control and cueing on figure of merit were used. For a score of reasons the instantiation to the full the policy of maintenance with respect to state of systems within classic understanding for wide-ranging total of guided aircraft missiles as exploitation objects is impossible. Therefore for the instantiation of operation and maintenance of guided aircraft missiles with respect to state a new method can be used.This method was named " the method of the utmost complete usage of the real technical state of a guided aircraft missile while decision making in situ.

The instantiation of such method assumes usage of the multialternative checking circuit: effective — qualified fit — defect, based on the assessment of the technical state of the guided aircraft missile. For this purpose a new notion J was introduced that was nameda combat potential of the guided aircraft missile. J was determined by the total of tactical and technical missile characteristics j₁, j₂, j₃,..., j_k, providing its dissemination against specific crowd of goals, under the crowd desired condition of the tactical employment. Thereafter the fault if only one any of parameters X ( of the missile as the unit under test), defined in the space Q^R(X) of the system of linearly independent permissible limits, spasmodically lower on 100% the whole combat potential of the missile notwithstanding that a limited stamina of the unit under test Q^R(X) stayed behind.

Let us introduce a new mathematical object F — the functional-hardware potential (FHP). F identically specified by the crowd of subsystem duties of the guided aircraft missile. These duties strictly reflect the sequence and the quality of carrying out missions of the guided aircraft missile. The above mentioned crowd was in practice implemented on missile board by defined total of on all fours of instrument room decisions. Let us analyze the class of the guided aircraft air-to-air missile subsystems as persistent dynamic systems in terms of the guidance system. The engineering philosophy of development of the standard guidance system is based on the condition for miss minimum R_h while target homing. Then the design of high performance securing control process is possible on the ground of construction and analysis of mathematical modelsfor the guided aircraft missile as the unit under test using figure of merit W of the true miss.

Hence it becomes possible the presentation of a monotonic association between the technical state of S missile subsystems in the figure of high merit space and the combat potential J of the guided aircraft missile. The usage of the multialternative checking circuit: effective — qualified fit — defect does possible the made-up enhancement of Ώ`(X) — the area of observed intermediate technical states of the unit under test. This area features the guided aircraft missile stamina which forms the requirement for the instantiation of advanced the guided aircraft missile performance strategy inside of the modern system of technical exploitation.

Paper is devoted to developing of pad elastic suspension designing methods for pad journal bearing with essentially new action mechanism. Developed bearing is tilting-pad journal bearing operating under oil starvation conditions so it is extremely economical. System state and dynamic stability provides working gap force closure due to tilting pads which are elastically suspended on the material MR insets. Elastically suspended pad journal bearings have the ability to high speeds operation and increased resource. Calculation to determine and optimize the performance of elastic- damping suspension bushings is one of the most important stages of a complex technique of designing these bearings. Elastic suspension material allows pads to turn in the radial direction and to make the best working position. The paper presents the developed mathematical models of working gap — pad — elastic suspension system behavior, created with the purpose of determining the elastic suspension material characteristics, providing bearing optimum operating mode. The mathematical method is developed for determining center of the resultant forces in system working air gap — pad — elastic suspension. The method uses the optimal load value, which is determined by the program calculating results (characteristics of lubricant flow: total pressure distribution along the working gap length). Obtained dependences can be used to optimize the pad geometry and MR characteristics (wire diameter, etc.). Model for determining the optimum characteristics of the elastic suspension material (wire diameter, wire material MR density, load value, etc.) is based on the V.N. Buzitskiy and A.A. Troynikov material MR model and takes into account the design and operational features of the bearing. Optimization of these elastic suspension parameters is a condition of failure bearing operation in mind its low oil flow, supernarrow working gap (from 1-5 mсm) and a mechanism for pad tilting on elastic suspension. On the CFD results basis (oil-flow characteristics, ANSYS CFX) using the developed mathematical models it is possible to determine the required MR characteristics and also the parameters of the MR layer behavior under loading by the pressure in the working gap (ANSYS). Two-dimensional contact model was created for system pad — elastic suspension to determine the strain value. Calculating of material MR stiffness C is based on experimental data. Expression for C is used to set the MR Youngs modulus dependence from the magnitude of MR deformation. The hydrodynamic load W = f (φ) is determined by the results of the flow characteristics calculation (pressure distribution along the length of the working gap). Distributions of radial and equivalent stresses, radial and total displacements (deformation) in the system were obtained. Through construction and investigation of the developed models characteristics the possibility of this technique is extended to the elastic suspension parameters optimization by various criteria. Established complex model of the system working gap — pad — elastic suspension allows to solve problems of associated hydrodynamics, thermal analysis and deformation.

The paper presents semi-empirical and theoretical methods for brush seal leakage performance evaluation. It also describes suggested simplified approach based on the reevaluation of the effective clearance value for the sel ected generic brush seal of the design that corresponds most closely to the studied seal. Besides,the paper describes the simplified CFD-based porous medium model of the gas flow in brush seal developed in free and open source software, allowing performing the analysis in automated mode. The proposed simplified approaches are practical tools for evaluating leakage performance of brush seals without excessive time consumption. The preliminary design approach is based on the experimental data on leakage and blow-down behavior of six different configurations of brush seals. Experimental dataset is supplemented by the predicted calculated relationships for the brush pack compression. In spite of the limited set of the generic brush seals, the proposed preliminary design approach demonstrates a satisfactory agreement with the data for other brush seals taken from free access literature. The simplified CFD model is used to obtain pressure and flow fields in the brush seal channels. The application of the proposed approaches is illustrated using the data on the original design of brush seals developed at IVCHENKO- PROGRESS SE. The results show that a reasonable agreement can be achieved for the brush seals of that differ significantly fr om the designs of brush seals forming the basis of suggested approaches.

The paper considers problems of selection of the joint of ceramic shell and metal bulkhead of nose antenna dome, that meets the condition of available analog design rethink minimum, while maintaining all the geometric dimensions of parts used. We show by calculations the process of design styles selection with further optimization of their geometrical parameters. Application of the other constructional materials used for a given assembly is inadmissible on radio engineering and mass considerations. The results of computational optimization have been proved by the field tests of prototype items.

Based on the results of the prototype air-to-air missile tests, the frame design has been corrected, and the maximum stresses in the ceramic shell of the fairing with the frame of corrected design have been determined. Application of modified frame together with the maintenance of air-tightness contributes to the product weight saving and stresses reduction in the ceramic shell, as well as improves the production effectiveness of the frame assembly.

At operation mode of 110 s duration maximum temperature of the inner attachment frame heating changes from 220 °C (at the shell edge) to 340 °C (in a thin part of the frame). The maximum tensile stress in the shell caused by the temperature is 15 MPa, and the maximum tensile stress caused by the force factors does not exceed 9.5 MPa. The total maximum tensile stress in the shell at the end of the operation mode is 17.3 MPa in the circumferential direction and it is 15.85 MPa in the meridian direction.

With all abovementioned taken into account, we recommend to accept the proposed design of the frame and to refine it with the aim to determine optimal geometrical parameters and develop the technological process which will make it possible to meet all the requirements of the frame design specification.

After analysis of frame design advantages and disadvantages, we plan to develop a new variant of the frame design with the following positive characteristics of previously analyzed structures, which will provide:

• air-tightness regardless of the frame assembly conditions;

• minimum stress in the ceramic shell;

• minimum frame weight;

• fewer parts to be assembled with the use of the sealant, simplified frame assembly.

A comparative analysis of the unidirectional action mechanisms, which include the irreversible motion clutch, self-braking electromagnetic clutch and transmission, is carried out. The purpose of the analysis is creation of an electromechanical drive of unidirectional action, which has the reduced power consumption, compact design, lower weight and improved dynamic power characteristics, for its use in control systems of aircraft of the new generation, rotation of power plants, jet rudders, flaps of water and oil regulators, and etc.

According to the presented theoretical research the following conclusionscan be made:

1. Currently, the unidirectional action mechanisms, which include the irreversible motion clutch, self-braking electromagnetic clutch and mechanical transmission, are very promising for the designers of electromechanical actuators for the new generation aircraft

2. The classification of the unidirectional action mechanisms depending on the relative position of the motor axes and the output link is presented; recommendations for their best layout are given.

3. As a comparative analysis of such class mechanisms in the technical literature is absent there was carried out a comparative analysis of the mechanisms of unidirectional action by mass, size, power consumption and dynamic characteristics.

4. It was found, that the mass of the cylindrical gear with braking profile is commensurate with the weight of worm and spiroidgears with load torque up to 50 Nm, then the mass begins to rapidly increase. In these mechanisms with more mass the load torque is more, that is why to put such mechanisms in multi-stage gear closer to the motor is advisable.

5. When the load torque is up to 15Hm all mechanisms have the same moment of inertia. With the load from 20Nm and higher the inertia moment of irreversible movement clutch significantly increases, and, besides, the inertia moments of the electromagnetic clutch, spiroid and worm gear is practically constant. With the load of 60 Nmthe inertia momentof braking profilecylindrical gear transmission is significantly increased. For highly dynamical air drive it is advisable to use the electromagnetic clutch with smallest moment of inertia.

6. To reduce the weight and size parameters the unidirectional action mechanisms should be placed between the motor and the first gear stage, or directly in this stage.

7. The greatest efficiency will have an one-way mechanisms of action, inwhich, depending on the functional tasks, the gearbox and clutch or irreversible motion, or electromagnetic friction clutch, are combined.

This work investigates the possibility of strengthening of production tooling, used for aircraft engine parts production, and cutting tools by diamond-like coating. We have developed diamond-like thin-film coating technique for the finishing plasma strengthening based on disintegration of liquid technologic agents vapors, ins erted into an arc plasma gun followed by plasma- chemical reactions and formation of product coating. We used metal-organic and organic compounds in liquid state as parent substances for providing diamond-like coating based on silicon oxycarbonitride. We choose argon as an orifice gas, used in arc plasma gun.

The applied coating duplicates the substrate profile, and strengthening coating herewith is formed as an optically transparent film.

We investigated comparison characteristics of adhesive properties of diamond-like coating and titanium nitride coating applied by ion-plasma sputtering in vacuum. As a result of our investigations we found that the val ue of critical load initiating first chips and delaminations equals to 35N for titanium nitride coating and 65N for coatings applied by diamond-like thin-film coating technique.

We found that being a dielectric itself the applied diamond-like coating forms a film barrier preventing stiffening of conjugated surfaces. Besides, this coating possesses higher corrosion stability and heat resistance. This fact is confirmed by numerous long-term atmospheric corrosion tests of samples under temperatures up to 1000 — 2000 °C.

For friction and wearout tests we used samples 38 mm in diameter and 12 mm thick made of steel P6M5 coated by means of diamond-like thin-film coating and titanium nitride coating applied by ion-plasma sputtering in vacuum techniques. As a counterbody we selected samples made of steel SHH15 heat-treated to the hardness of HCR 63. Contact conditions were as follows — rolling friction with 20% sliding with lubrication. The tests were carried out at 1000 min-1 rotation frequency and 1650 N load. For comparison, we also used a sample made of heat-treated steel P6M5.

In the course of the experiment, the values of friction torque and mass wear of the samples were registered, and friction coefficients and wear rate were calculated.

Based on tests results we determined that the friction coefficient of the surfaces with diamond-like coating is two times smaller, compared to reference surface of steel P6M5 and 17% less than for the surface with titanium nitride coating applied in vacuum. The wear rate is reduced by 46% and 16% correspondingly.

Due to the obtained data we came to a conclusion of the expediency of application of the developed technique for strengthening of production tooling and cutting tools in various industries, and for aircraft gas- turbine engines parts manufacturing in particular. The developed technique was tested at the number of enterprises, and its application showed 3-6 times increase in production tooling and cutting tools durability.

The development of new technologies and equipment for laser processing of aircraft materials during the creation of aviation and space technology products of new generation is an urgent task. The development of new laser technologies requires the creation of highly technological equipment, integrating mechanics, electronics, automation and information technology. The control system is an integral part of the machine with intelligent control of their functional movements. Application software, which includes the qualitatively new modules, ensures the implementation of the control system of laser processing unit, and allows to create equipment for new processes. Speed laser processing and cutting machines are widely used in the creation of elements and devices for aviation and space technology.

Article is devoted to the development of application software to manage the new process of laser controlled thermosplitting, as well as technological processes for laser cutting and processing of materials in the new industrial equipment. The basis of the control system is a laser processing equipment software LaserCNC designed for different processes such as laser-controlled thermosplitting.

The enterprise Lasers and equipment TM has created a new industrial plant MLP1-1060/355 for the implementation of a new precision separation process of plates made from brittle materials based on the use of a new method of laser controllable thermal cleavage, which uses two lasers: one is to separate the plates, the other one is for the initial application defect for crack initiation.

Developed software LaserCNC for control system of the laser technological complex MLP1-1060/355 is possible to implement the innovative industrial equipment for the new LСT technology. The LaserCNC software supports different laser processing equipment and, also, can be used for the new multi-axis kinematic systems.

The control program for systems with more complex kinematic scheme is DopOs1, and there is DopOs2 for multi axis machines with additional kinematic axes. The basis of the control system is the laser processing equipment software LaserCNC, DopOs1, DopOs2 designed for different processes, such as laser-controlled thermosplitting, laser cutting and material processing. This software is designed for industrial laser-processing equipment with various kinematic schemes.

Creation of systems with more complex kinematic scheme will provide a significant improvement in dynamic performance and accuracy required in processes of laser cutting and material processing without cost increase due to the use of more powerful drive. The use of modeling program DopOs2 will make another step in the design of new industrial laser processing unit with additional kinematic axes.

The paper presents the results of flying vehicles high-ratethin-walled elements surfaces hardening by surface plastic deformation.

Investigations of mechanical-and-physical properties of metal surface layerat hardening by surface deformation were carried out.

The depth of hardened layer under optimal surface processing rates by hardening with surface plastic deformation.

We establish the relationship between quality characteristic of surface layer with endurance and durability of machine component parts, and performed their quantitative assessment.

Investigations were carried out and positive results were obtained, which allow manufacturing application methods of hardening by plastic deformation for a number of critical parts. These investigations included pilot parts hardening under different modes in laboratory conditions with bench testing and machines with pulsators to compare these parts endurances, as well as under manufacturing conditions with carrying out appropriate comparison performance test ofendurance.

When solving rotor dynamic tasks of rotating machines including aviation gas turbine engines, calculation accuracy of the system elements is highly important. One of the most widespread elements are roll bearings. In the first place stiffness characteristics of roll bearings are determined by the contact phenomena appearing between rolling elements and rings. As a whole the other elements of the roll bearing (such as inner and outer rings) influence stiffness no more than 5%. Interaction between the roll and the ring may be considered as linear contact between two cylindrical solids with parallel axes. For the first time solution for this task was obtained in the Lunberg- Sjövall paper [1]. However, equation in [1] has implicit dependence of outer force on appeared displacement and so cannot be used for tasks of stiffness characteristics determination. In Harris work [2] empirical Palmgren equation [2] was used to solve this task:

,

where — contact force, — contact length, —displacement in contact line.

To verify results in paper [2], the task of contact interaction between a half cylinder and a plane was solved using FEM. It was obtained that there was significant divergence (up to 30%) between solutions obtained using Palmgen’s equation and FEM. It was found out that Palmgren’s equation was correct with error of 5% for lengths of contact lines from 10 to 20 mm. According to the results of the FEM solution, the new dependence for linear contact tasks was proposed. It gives no more than 4% results’ divergence in the contact lengths range from 0,5 to 50 mm.

.

The obtained dependence may be used to simulate stiffness characteristics of the roll bearings and to describe linear contact in tasks of contact interaction mechanics.

The main purpose of the materials research in the aerospace industry is to improve the structure reliability while reducing the weight. Heat-resistance titanium alloys with a density of about 4,5 g/cm3 can facilitate engine weight, but titanium and its alloys have polymorphism, which limits the application temperature of the alloys in 0.6 times then polymorphous transformation temperature. A heat resistant titanium alloys can continuously operate at temperature of 500 550 °С, depending on the chemical composition of the alloy, the process conditions of heat and deformation processing. Discs and the compressor blades made of titanium alloys where used in aircraft gas turbine engines. Further reduce of weight of the engine can be achieved by replacing the classical design of the compressor stage with removable blades by blisks.

This article examines the impact of different modes of annealing on structure, phase composition and mechanical properties of the samples of punching blisks. Raising the temperature in the first heating stage at ten degrees Celsius and lowering the heating temperature in the second step of annealing from 590 to 550 °C reduces the volume fraction of primary á -phase and 50% in batch annealing to 2025% reduction in the thickness of the secondary plates á — phase and increase the length of the interphase boundaries. Using transmission electron microscopy it was found that when adjusted mode annealing at interphase boundaries silicide particles are not detected. By increasing the efficiency of solid- solution and precipitation hardening is provided by increasing: the strength of 120 MPa at room temperature and 70110 MPa at elevated; rupture strength at 500 °C at 75 MPa; low cycle fatigue at 20 °C for 20 MPa high cycle fatigue at 20 °C at 30 MPa. At the same time, due to changes in the structure and phase composition, sensitivity to stress concentrators during the impact test is reduced.

The necessity to increase weight effectiveness of new aviation structures encourages permanent structures modification and development implementing new materials, including composites. Some elements of these airplane structures can be subjected to impulsive loading. Thus, the investigations of dynamic deformation of shells and plates made of composite materials loaded by the impulse pressure are important. As long as the great majority of structures was made in the past of aluminum alloys, it is rather interesting to compare their deformations with those of new stuctures made of composite materials under the same loading to identify the differences and advantages.

The goal of the presented article is numerical and experimental investigations of two plates made of aluminum and composite materials, loaded by equal impulse pressure with maximum amplitude of 6-7 Bar order and 0.3-0.4 ms duration. The composite structure is made of three layers: two outer plates of fiber- reinforced polymer, linked by honeycomb stuff.

We realized numerical simulation is performed on the basis of explicit integration schemes of the equations of motion in time domain and finite element sampling in space in LS-DYNA application software package. Eight-node of Sollid type and four-node finite elements of Shell type are used for sampling. The problem is solved in the Lagrangian formulation. We used Chang- Chang orthotropiс model with possible brittle fracture for composite material. Тhe accuracy of numerical solution is confirmed by the precision of calculation data and experimental results.

Comparative analysis shows that numerical and eхpеrimеntal сurvеs of deformations in several sеlесted points are close to each other. Aluminum and сomposite platеs of diffеrent sizеs are deformed by thе samе dynamiс lоad up to maximum levеls of dеformations lеss than 0.2%. Numеriсal and eхperimental rеsults show that aftеr impulsе loading removal thе platеs keep on their oscillatory motion with the lowest fundamental frequencies. In the еxperimеnts thе period of osсillation damping is about 180-200 ms for aluminum platе and about 40-50 ms for сompositе plate. The three-layer compositе plate demonstrates four times higher damping effect compared to aluminum plate. From the above said we can conclude that numerical calculations of dynamic deformation of composite plate based on Chang-Chang orthotropiс model allow obtaining acceptable results. Further improvement of modeling quality can be achieved by more detailed description of honeycomb stuff.

The article is devoted to the issue of the dayconcerning calculation of the stress state of the overlapping glue jointin two-dimensional formulation. The purpose of the work is to solve the problem of glue jointstress state of symmetrical rectangular plates with different width under longitudinal load. For the first time we obtained the solution of the problem in analytical form. To build the solution we used simplified Volkersentwo-dimensional model, according to which connecting layers are considered to be rigid in transverse direction. The hypothesis of the plates rigidness in transverse direction was implemented earlier for creation refined theory of beams. The author developed this theory for connection of two plates. The outer layers work only in tension and compression in the longitudinal direction and shift. The adhesive layer is a shear layer. The stresses assumed to be constant over the thickness of the glued layers. The stresses in the adhesive layer are proportional to the difference between the displacements of the outer layers. Due to joint symmetry and imposed stiffness layers hypothesis, there is no movements of layers in the transverse direction. In the gluing area, the stress state is described by a system of two, second order partial differential equations with respect to the longitudinal movements of the outer layers. Outside the gluing area the movements of a wider layer are described by a second order partial differential equations with respect to the longitudinal displacements. Solution of the equations is build as a series expansion in the eigen functions. Eigen values are found fr om homogeneous boundary conditions at the lateral sides of the connection and the conjugation conditions on the lim it of the bonding domain and the outer edge of wider plates. The corresponding eigen functions are not orthogonal. Coefficients of the series are found from the orthogonality condition of the residual vector of boundary conditions to the eigenfunctionsvectors.

The model problem is solved, and the obtained data was compared with calculations based on implementation of the finite elements method. Calculations have shown that the accuracy of the method is sufficient for designing such connections.

Nanocomposites of different concentration of cadmium sulfides nanoparticles in polyacrylamide shell were obtained as a result of interaction of cadmium nitrate with acrylamide and sulphonating agent (thiourea) in certain stoichiometric ratios and subsequent thermal polymerization of the reaction mixture. The samples were studied using methods of elemental analysis, X-ray diffraction, scanning electron microscopy, optical and fluorescent spectroscopy. The size and distribution of sulfide nanoparticles in polymer were defined. X-ray diffraction analysis of nanocomposites obtained revealed the following reflections at 2Θ: 26.35; 43.85; 51.85 with the values of d equal to 3.357; 2.085; 1.756, characteristic for hexagonal cadmium sulfide with wurtzite structure. Intense broadband at 2Θ = 21 belongs to polyacrylamide. It is characteristic that the intensity of CdS-containing phase peaks decreases with the increase of cadmium sulfide concentration in the composite with a simultaneous increase in the interplanar spacing. It proves nanocrystals binding with a polymer matrix. The average crystallite size (nanoparticles), corresponding to the sizes of coherent scattering of X-rays ranges, were evaluated at the width of the reflections in the diffraction patterns according to Debye-Scherrer formula. Particles size varies within a narrow range of values and ranges from 4 to 15 nm with increasing of the CdS share in the composite. Increasing of the CdS concentration reduces the particles size of the CdS crystallites.

Submicroscopic image of nanocomposites containing different concentrations of cadmium sulfide nanocrystals indicates a noticeable effect of nanoparticles concentration on the surface morphology. That is, increase in the cadmium sulphide concentration leads to an increase in the compactness and structuredness of the composite.

It was shown that the absorption spectra has characteristic features of the spectra of composite systems with nanosized semiconductor crystals, which show «defective» luminescence with maximum in the range of 550–570 nm.

This article examines the new type of the aircrafts — electrically drivenaircrafts (electro-aircrafts).Presented two trends of development of present-day electro- aircrafts, which use electric energy for mission operations. The first trend lies in meansimplementation ofsolar radiation energy by converting it on-board into electric energy by means of photoelectric converters allocatedon the surface of the aircraft. The second trendisimplementation of the electric energy stored on boardin batteries. Each of the trends include either manned or unmanned (UAV) prototypes of electro- aircrafts. From the first trend, so-called solar UAVs (or SUAVs) have widely spread in the forms of sport model aircraftsand industrial designs. The paper presents NASA achievements in the field of electric UAVs, designed in the course of its twenty-year ecology program ERAST, which at present time has expired. It also observes NASAs challenges for the development of such aircrafts in the future. The paper examines design features of SUAV Helios HP01. It is noted that even at a low value of efficiency of solar batteries SUAV can be competitive with satellites for monitoring the Earth surface. In addition, the paper notes the works with UVA carried out in MAI. The layout of the first Swiss man- carrying aircraft driven by solar energy, which was in flight within 24 hours is presented.

The second trend of electro-aircrafts development was presented for the first time at Berlin ILA-2014 exhibition by two existing prototypes. The main attention there was focusedon the electro-aircraft E-Fan, presented by the Airbusgroup. This flying vehicle can be considered as a prospective for ecological transport aviation. The paper presents its technical characteristics, and marks its design features based on composite materials, Li-Ion batteries, allocated in the wing, and original propelling screw.

Austrian firm Diamond Aircraft Industries involved in the development and creation of small aircraft prototypes presented its vision of on-board electric power source. At Berlin exhibition, it demonstrated its electro- aircraft LANDE with promising on-board battery, based on hydrogen fuel cells, similar to those used in the space industry. It is noted that such power systems were designed for UAV at MAI. Experts believe that in the future hydrogen energy sources will replace batteries for electric transport helicopters. Helicopter firm Agusta Westland designed, manufactured and tested with electric system and original structure of the flying wing type with two swivel propellers for vertical takeoff and landing.

Two micro UAVs, built by helicopter and aircraft schemes with electrical on-board systems-MIKADO and ALADIN correspondinglywere presented at Berlin exhibition. These flying vehicles have small mass and size, and are intended for near-field reconnaissance by one fighter. They were tested in Afghanistan for air reconnaissance, mapping and other military operations. At present, armies of Germany and NATO have a great number of such aircrafts.

All prototypes of electro-aircrafts presented are promising.

This work aims at the analysis of the rotary electromagnet structure through the research of electromagnetic field patterns by finite element method. Based on the results of the analysis we built a general equivalent circuit of a magnetic circuit, enunciated assumptions necessary to build non-linear mathematical model as well as electromagnet of rotary type design.

The analysis of numerical values streams obtained as a result of electromagnetic field modelling in a system with neutral electromagnets shows that in most cases stray and bulging fluxes appear to be commensurable with working fluxes, and, thus, they need to be taken into account when calculating the static and dynamic characteristics. In the electromagnet under consideration, the stray flux is almost constant during the movement of the armature. It allows supposition on stray conductivity uniformity.

On the basis of the analysis of electromagnetic fields patterns and assumptions, we build a non-linear the electromagnet mathematical model.

Research of electromagnet dynamic characteristics was carried out numericaly using non-linear model for ambient temperature vatiation of ± 50°С and power source voltage within the specified tolerance range.

Mathematical model obtained and developed software can be implemented to analyse currently available as well as newly designed electromagnets. The developed mathematical model takes into account the peculiarities of magnetic field configuration of each electromagnet design implementation.

Realization of reforms in the science-intensive business is accompanied by associating of existing enterprises and the creation of new ones. The rules of law, corporate and professional codes govern the corporate relations between enterprises. These relations are characterized by a certain degree. These relationships between enterprises are characterized by a certain degree of dependence on interaction with each other.

For the Russian aviation industry, the role of the state at an early stage of economic reforms seems of less importance compared to the times of the USSR existence as well as leading world economies.

Since the beginning of economic reforms, production volume of aircrafts has declined significantly. The quality of designs deteriorated, and for the most part represented an adaptation of previously designed products. Commercial aviation manufacturing became dependent upon constituent parts of foreign manufacturers. Defense technology designs achieved more recognition since they are related to the sixths generation.

Research and development activities significantly decreased, and breakthrough innovations were replaced by a step-by-step modernization. For research and development effort organization everybody stuck to the principle of so-called pulling-out. Taken together, it led to intellectual potential reduction, transition to works, related to designs of small batches of aviation products manufacturing, independent search for orders, and consumer goods design.

At the early stages of the reform of educational institutions, a significant outflow of personnel was noted. This outflow of personnel had a significant impact on training.

The conditions of admission to educational institutions and training have also changed: training on a commercial or budget basis and targeted training were offered. New forms and training technologies are used.

The number of graduates of educational institutions did not and do not meet the demands of the industry, academic and sectorial research institutes for the specialists of appropriate qualifications.

Over the past 10 years, the federal government adopted standards for training, the implementation was carried out only by scientists and teachers of high school.

State program of development of the aviation industry for the period from 2013 to 2025 with the subprogram 7 Aviation Science and Technology is aimed at solving problems arising in corporate relations in the period of reform.

Implementation of reforms in high technology business association is accompanied by existing businesses and creation of new ones. Corporate relations between companies are governed by rules of law, corporate and professional codes, characterized by a certain degree depending on the interaction with each other.

Role of the state for the Russian aviation industry in the early stages of economic reform is recognized lower compared to the time of the Soviet Union and in comparison with the worlds leading economies.

Since the beginning of economic reforms, production of aircraft has declined significantly. Quality development deteriorated and is a modification of previously developed products. Production of civil aviation has become dependent on the components of foreign origin. Developments in the field of military equipment reached more recognition, they are referred to the sixth generation. In the manufacture of preserved preference of domestic components.

Research and development activities significantly decreased, breakthrough innovations are replaced by a step by step modernization in the organization of remained the principle of pull. Overall, this resulted in a reduction of intellectual potential, the transition to work-related developments for the production of small batches of aviation products, independent search of orders and product development of national purpose.

In the early stages of the reform of the educational institutions there is a significant outflow of personnel, which affects the quality of training.

The conditions of admission to educational institutions and training have also changed: offer training on a commercial basis or budget, conditions targeted training, new forms of technology and training.

The number of graduates of educational institutions did not satisfied and do not meet the needs of industry, academic and industrial research institutes in frames appropriate qualifications.

Over the past 10 years, Federal State Standards for training of specialists were adopted, which implementation was carried out only by scientists and teachers of high school.

The State Program of development of the aviation industry for 2013-2025 with the Subprogram 7 Aviation Science and Technology is aimed at solving problems arising in corporate relations during the reform

This article deals with developing of knowledge about innovation management mechanism in the context of researching national innovation system. We studied key factors that influence innovation process at the enterprises. We tried to develop key factors that help people in their innovation activity and that stem the tide. It was in our interest to define key factors that help the process of knowledge diffusion at the enterprise. Motivation, as we know, is the key element of innovation management mechanism. Also we studied innovation culture and some notions of it.

First, we studied some researches in the sphere of social oriented mechanism of innovation management as Russians authors and foreign. We tried to define key factors — motivator for innovation process so among shareholders and management, so among liner managers and specialists. We made expert survey using structured questionnaire. The number of experts was more than forty. We define the expert as a specialist with working experience in the sphere of innovation activity in different area of economics passed special courses of innovation management and having MBA or PhD degree.

We found out that in very large companies the most important factor that prevents the innovation process is the lack of financial recourses, high cost of innovation, not enough government support. In amount of middle, small and little companies the most important factors are: support of management, composition of project team.

We point out that knowledge management play the key role in innovative process and the key factors helping knowledge diffusion are support of company management and affiliation to some innovation team.

Innovation culture is associated with such notions as management participation in innovative process, fair incentives system, atmosphere of trust and striving to personal growth of each member.

Introduction of cycle life support processes of science-intensive production into an aircraft industry is an actual problem at the enterprises of the aviation industry, becauseat the present time many enterprises of aircraft industry are faced to the problem to chose the integrated information system for increase of enterprise activity efficiency.

The basic principles for introduction of the integrated information management systems are defined by tasks and features of the aircraft equipment production, and, also,by specifics and development of modern methods and means. In this regard the main objective of introduction of information support into life cycle processes of aviation products is finding of an optimum combination of separate methods and means of information technologies.

During the solution of the mentioned economic problems the information technologies, which are directly aimed to support of productions at all stages of life cycle of the created science-intensive aviation production, have the great importance. During the analysis of vital processes of the science-intensive production of aircraft industry it is revealed that the full- scale introduction of the integrated information systems at various stages of life cycle of the aircraft equipment is capable to bring to the developers, producers and customers the following advantages: at the design stage of products, tests and operational development, and also technological preparation of a mass production — the possibility of reduction of duration and cost of the listed stages, possibility of flexible change of a design of products and technology of their production; possibility of formation of the virtual enterprises and flexible change of their structure for the purpose of ensuring low deficiency and purchase price of accessories; possibility of optimum planning of processes of maintenance, and also material support of their operation; possibility of the organization through imitating modeling and optimization of life cycle of the aircraft equipment since its earliest stages. Realization of the principles of information support of the science-intensive production is possible only in the case of complex introduction of information technologies at all stages of life cycle of production of aircraft industry.

Operation of air data system (ADS) assumes the application of atmosphere static and dynamic pressure measurement affecting an aircraft. At present the type of airspeed tube construction and a place of its mounting are selected through experimental research in wind tunnel [1, 2].

With computing technologies development, we got the possibility to predict gas pressure distribution over aircraft surface depending on an aircraft design, flight conditions and aircraft state vector [3, 4].

Pressure distribution determination over aircraft surface using up-to-date software is a relevant task, allowing advance in mathematical design of ADS and substantially reduce the extent of experimental examinations, providing high accuracy of measurements due to more detailed measuring process models.

Having calculated pressure distribution over aircraft surface with due account for change stipulated by deflection of incidence angle and gliding, velocity, Reynolds number, etc, one can optimize the point of airspeed tube mounting, as well as provide theoretical calculations of pressure measuring error. All this contributes to compensating this error programmatically with ADS onboard computer.

The goal of this work consists in obtaining static and dynamic characteristics of processes taking place in airspeed tube.

The provision of store release safety is one of the issues that arise during the development of aircraft store transportation and separation systems. In general the safe release refers to fulfilling the following requirements during the store flight in the vicinity of the carrier after its release from the launcher:

— elimination of the possibility of collision of the store with the airplane itself, its structural elements or stores, which are installed on the adjacent suspension points;

— retention of the store spatial stability on the initial segment of its flight path in the vicinity of the carrier;

— absence of thermodynamic impact from the exhaust jet of the store engine on the structure of the airplane itself,its propulsion system functioning and other stores.

It is possible to solve the problem of store separation safety at the stage of transportation system design and release method development. The appropriate solution method is based on the analysis of the results of modeling of store spatial movement in the vicinity of the airplane. At that the reliability of aerodynamic characteristics of the isolated store and flow, which is disturbed by the plane, is important for the simulation.

As a rule it is necessary to conduct wind tunnels tests to obtain aerodynamic characteristics. Reliable determination of safe release modes requires conducting a large number of tests at different angles of attack, sideslip angles and Mach numbers. The presence of folding aerodynamic elements of various configurations on the suspended stores increases the amount of required wind tunnel experiments dramatically. Thus, the determination of aerodynamic characteristics of isolated bodies and interference can take a long time. Conducting tests is generally not feasible at the early stages of system design under uncertainties of some design parameters. In this case it is expedient to use computer technologies, namely, computer engineering methods. Such methods provide reliable analytical support through the use of software suites, which are focused on computational analysis of gas and fluid dynamics (CFD).

The software system, which performs CFD analysis and processes the virtual experiment results, was applied to the analysis of an external fuel tank. Such fuel tanks are widely used on various types of aircraft. Virtual simulations were conducted in a wide range of angles of attack, angles of sideslip and Mach numbers. The results were processed by using a multidimensional (multivariate) interpolation function. Such function allows the research team to calculate the values of the aerodynamic coefficients for the entire range of angles of attack, angles of sideslip and Mach numbers. The relevance of the obtained aerodynamic characteristics was confirmed by the comparison with the available data of the wind tunnel tests. The comparison showed good consistency between the calculation by CFD method and full-scale experiment. Slight differences are most likely associated with the following circumstances:

— ascaled-down model of the fuel tank was tested in the wind tunnel,

— presence of a methodological error in the algorithm, which was used for solving the equations,

— presence of methodological error in the algorithm of gas flow simulation,

— the three-dimensional solid model was simplified for the analysis.

The described method allows the designers to determine the aerodynamic characteristics of the stores at early development stages without conducting the wind tunnel experiments.This can reduce the cost of wind tunnel tests and time, which is required for optimization of the store aerodynamic configuration, significantly.

Application of the proposed multidimensional (multivariate) interpolation allows the designers to calculate the coefficient values for any angles of attack, angles of sideslip, Mach numbers, geometrical configurations and other initial data within a pre-defined range.

The application of the described method allows the researchers to calculate the aerodynamic characteristics of both newly designed and existing aircraft in a wider range.

All of the mentioned above confirms the efficiency of computer engineering methods, especially at the early stages of aircraft design

The purpose of this work is to validate model of airflow over passenger aircraft aerodynamic using the ANSYS FLUENT software package (license number: 00632255).

The paper examines two models of a flown around object: LL 6945 fuselage model [1] and TU-154B passenger aircraft [2]. All investigations were performed using the following parameters: angles of attack, α = 0...20°, Mach numbers: 0,2 < M < 0,8, Reynolds number: 7·10⁶ < Re < 28·10⁷. The flow in the vicinity of the examined models, LL 6945 fuselage and TU-154B was assumed to be turbulent.

Calculations of the flow parameters and aerodynamic characteristics were made using the following models of turbulence: Spalart-Allmaras model, «SST k- ω», «k- e ».

Calculation results obtained with ANSYS FLUENT were compared with experimental data.

It was found that calculation results using the «SST k- ω » model match in the best way with experimental data and describe the variations of the aerodynamic coefficients with respect to the angle of attack in comparison with other models of turbulence [1, 2]. Constants which give the best convergence of calculated and experimental data are determined for the «SST k- ω » model. The results obtained will allow application of this flow model to optimize pitot-static systems.

Development of aviation technology requires the use of a variety of measuring systems to determine the motion and positioning variables of the aircrafts in the space. Requirements for the reduction of weight and dimensions of such systems are the first since the objects themselves demonstrate a tendency to reduce their weight and size characteristics.

Availability of inertial measurements units (IMU) allows to create tiny strapdown inertial navigation system (SINS), which have small weight and dimensions, and can be used in areas such as attitude control system, radar, avionics, data transmission system, navigation, etc. The work of these systems is accompanied by a variety of errors and failures, which leads to the urgency of early detection and recognition of their errors. This difficult problem can be solved on the basis of different approaches and with varying degrees of success.

The article describes one of the possible ways to solve this problem, such as using combination of two methods like a parity space approach and the discrete wavelet transform in the redundant system with additional IMU. Such combination allows to detect three faulty sensors and to isolate each of them out from the system during the system operation. The one of the positive advantage of this algorithm that it is not demanded to know the dynamic of the system model, except orientation of the sensors in the measurement system organized with redundant IMU. The simulation of the successful operation of this algorithm with two types of sensor faults, such as bias and stuck, are demonstrated in this work.

The article describes optical micro electromechanical(MOEM) angular velocity transducer, representing a solid-state device having vibration console, with the ability to deviate. This transducer allows determine magnitude and direction of angular velocity by using Coriolis force effect and optical information reading. Coriolis force is acting on the piezoceramic sensitivity element during rotation of the MOEM transducer, causing its deviation over the axis, positioned on vibration plane and directed perpendicular to the plane along the longitudinal axis of the console.

Material and dimensions of sensitive piezoelectric element (SE) should be such that the resulting deflection of its parts does not exceed hundreds of nanometers. The highly sensitive optoelectronic circuit for information retrieval about angular velocity with the purpose of obtaining more accurate information should be herewith based on optical tunneling effect (OTE), allowing fixation of a small measured displacements.

The structural model of uniaxial MOEM angular velocity transducer is described. The impact of the Coriolis force on SE is calculated depending on various design parameters and piezoelectric materials. For a SE in the form of a plate deflection has no strong dependence upon the temperature, and vibration element longitudinal oscillations amplitude linearly depends on the coordinate along the longitudinal axis and exciting AC voltage.

Simulation results showed that numerical values of the deviations can be obtained within the range of several hundred nanometers, which corresponds to the value of the optical radiation wavelength. Thus, angular velocity transducer can be designed on the basis of optical tunneling effect, which provides greater sensitivity when measuring small angular velocity.

The above characteristics analysis shows that the material with the brand ZTC-19 having larger piezoelectric module and higher stability within the working range of temperatures and pressures, can be used in MOEM angular velocity transducer based on OTE.

We examined characteristics of the transducer, using SE with additional mass to improve its sensitivity. With structural dimensions decreasing it is necessary herewith to increase the value of additional mass to provide greater sensitivity for measuring small angular velocities.

With the same sizes and materials introduction of additional mass m = 5·10⁶ kg, SE end deviation under the influence of the measured angular velocity will be 1,5 µ m, which is hundred times greater than in the case without the additional mass m. It also provides the possibility of organizing an optical information reading by means of optical tunneling.

The problem of surveillance and search of highly dynamic objects based on the image processing and analysis of video in real-time is considered in this paper. By highly dynamic objects, we mean such objects, which interframe shift exceeds their correlation radius. The realtime operation condition depends on the current task specifics, and can be expressed as a required dynamic objects computational efficiency (the amount of elementary operations) of the search algorithm.

As an example we examined search correlation algorithm MAD and search algorithm SIFT based on descriptors. We carried out studies of these algorithms and specify conditions under which these algorithms would fail to provide realtime condition fulfillment. The highly dynamic objects search algorithm ensuring realtime operation under these conditions is presenter in the paper. The advantage of the proposed algorithm lies in significant reduction of elementary operations due to required object descriptions presentation in the form of star pattern of characteristic points.

It is shown that proposed algorithm has higher computational efficiency than correlation algorithms and algorithms using descriptors.

We carried out experiments of real time objects tracking, and determination of the blade edge position of helicopter propeller. The experimental results show the efficiency and effectiveness of proposed algorithm application for tracking and tracing problems.

The article reviews a problem of layout design automation for aerospace engineering. It is shown that the complexity of this process is stipulated by the difficulty of information representation concerning geometric shape of an object being arranged in the computer.

The article presents the main types of geometric models describing the shape of objects being arranged, their classification and application areas. We describe an automated layout problem as a problem of placement of a certain group of objects of specified sizes and shapes within a confined space. The task of automated layout is formulated as mathematical programming optimization problem of for maximum allocation tightness of objects being arranged. We present optimization criterion of this process that is space factor of a space being arranged by objects allocated inside it, as well as limitations of this optimization problem.

The article also describes well-known methods of automated layout. Powered by article analysis shows that they are implemented only for objects of simple geometric shapes — primitives. Proposed a method for automated layout, based on mathematical formalism of normal equations. The concept of objects normal equations was introduced by V.L. Rvachev. This article describes the original methods of normal equations building for both flat and tridimensional objects. It is shown that application of normal equations of objects being arranged allows you to create intelligent layout algorithms, reducing in terms of geometry the position problem of allocation of objects being arranged to the metric one. The paper shows additional options of objects normal equations for automated layout problems, such as equidistant object building. We display the problems that should be solved for a wide use of normal equations apparatus in automated layout practice.

The subject of research is a lunar base with nuclear power station (NPS) containing one or more thermoemission modules (TM). The purpose of current work is determination of the radiation situation in the neighborhood of the lunar NPS. The task of calculating radiation shielding (RS) in the neighborhood of the lunar NPS with one or four TM was solved for the implementation of this goal. The thermoemission nuclear power plant (NPP) with electric power of 50 kW was considered as TM.

Calculation of radiation functionals was produced using the codes which allow to obtain a solution to the transport equation in two-dimensional (RAPID-KONT) and three-dimensional (MCNP) geometries. These codes implement two different methods of solution: a deterministic method and the Monte-Carlo method. Optimum profiles of RS layers were determined by using contribution currents and methods of optimal control theory (the Pontryagin maximum principle). The induced activity of the lunar soil and the fission products activity were calculated with the KAMOD-K code.

The following variant of location of the reactor and RS in mine was sel ected as a result of designer research. The depth of the mine is to 3.7 m and the diameter is 1.41 m.

The optimal mutual location of four NPS, substations and lunar base was received for sel ected variant of the TM placement. The total mass of communications between service module and each of the 6 TMs (two in a reserve) is ~2.1 t, the mass of communications between service module and the lunar base is ~0.4 t, the distance from service module to the lunar base is ~150 m.

The radiation situation in the neighborhood of stopped NPS is mainly determined by the doze of gamma-radiation fr om the activation of the lunar soil. The dose of gamma radiation fr om the remaining activity of stopped reactor and the induced activation of the soil is reduced in 5 times by filling the mine by lunar soil.

The obtained results can be used in manned lunar base designing. The number and capacity of each TM is determined by the needs of the consumers.

Thus, it can be stated that the radiation situation in the neighborhood of NPS allows the personnel work at the lunar base and service module when TM is located in the mine.

The issue concerning the design of the stabilization system of cable tensionis being solved in the given work.

On the basis of the conducted study and review of the similar systems the structure of the stabilization system of cable tension was suggested. This system is electromechanical two mass elastic flexible system with the interval parameters. The length of the cable and the mass of the apparatus under test are considered as theinterval parameters. For the given system the mathematical model was developed and the control and disturbance transfer functions were obtained.

On the basis of the application of the robust approach to the synthesis of the system under consideration the methodology of the definition of the robust regulator adjustment ensuring the tolerable quality coefficient of the system operation at any values of its interval parameters was developed. The criterion of the maximum stability degree was used as one of the criteria of the robust regulator synthesis being the most relevant for the development of the system with unstable parameters. To apply the mentioned above criterion it is suggested to use the sufficient conditions of the specified stability degree of the interval system by the limitation on its oscillation and accuracy. The specified conditions can be represented as the inequality systems binding the interval coefficients of the characteristic polynomial and its coefficient quality indices (stability and oscillation parameters).

From the obtained interval values those parameters are only chosen which ensure the maximum stability degree and minimum oscillation. To check and test the obtained results the areas of root locus of the interval characteristic polynomial are being constructed.

The simulation modeling of the synthesized control system of cable tension in the Simulink package is carried out to assess the operability of the developed methodology of the synthesis of the robust regulator. The obtained transient characteristics under the worst modes of operations confirm the accuracy and validity of the selected solutions.

Software was developed to put into practice the methodology of the synthesis of the robust regulator. This software allowsin on-line mode to make calculations of the desired parameters of the robust regulator with maximum stability degree of the system by its oscillation degree limitation.

The results of the scientific investigation obtained in the given article can be used at aerospace industrial plants.

In the context of problem of all-weather location of distant space objects four original detection schemes of far IR radiation, approaching by their sensibility to the level, allowing to use them in photon counting regime, are described.The first scheme is actually an upgrade of technique of the superheterodyne transfer of quantum fr om far IR to the visible range, wh ere photon counting is produced by a usual photomultiplier or APD (avalanche photodiode). The design of Up-converter considered assumes that nonlinear crystal-mixer is placed within the resonator of the single laser block. The second scheme of registration of far IR is based on directly biased light-emitting diode when the current is yet insufficient for radiation generation. Realized experiments allowed to observe photo response of such a system on radiation with energy below red edge of internal photoelectric effect. Two last schemes exploit cryogenics: in the first one far IR photon counting is assumed to perform using a number of desorbed helium atoms which are registered, and in the second — a number of unpaired spins in antiferromagnetic, the magnetic moment of which is measured by a SQUID. In both ways the effect has a nonbolometric character that provides speed sufficient for an application of such schemes in pulsed location systems. In the first case the nonbolometrisity of the effect is the consequence of stability of temperature value, corresponding to the first order phase transition, when all incoming power is spent not at material heating but exclusively on transforming the matter from one aggregation state into another. In the second case nonbolometrisity is achieved due to direct resonance action of far IR radiation on antiferromagnetic, resulting in flip-flop of spins and appearing of magnetic response registered by the quantum interferometer.

In conclusion two possible registration schemes of near space small objects that do not use the location in ordinary sense are being considered: photographing in 10 µm range using a CdHgTe matrix and pulsed CO2 laser with transverse pumping (TEA ) as a megawatt flash; space gravi-exploration using the system SQUID-magnetostrictor, offered initially as an ultrasensitive gravitational wave detector.

Prediction analysis of foreign and domestic studies shows that turbofan with high BPR and high cycle parameters using aviation kerosene will not provide meeting to environmental requirements for new aircraft already after 2035. And only hybrid gas-turbine engine (HGTE) using liquefied natural gas (LNG) allows developing aircrafts, meeting stringent emission standards up to 2045.

The purpose of work is studies of available foreign and domestic research experience and generation rational concepts of HGTE using various fuel types, including gas and cryogenic fuels.

According to purposes of the work following activities are conducted:

— results of a foreign forecast of HGTE, conducted by Boeing and NASA, are analyzed;

— perspective concepts of foreign power plants (PP) for aircrafts 2030 with HGTE having potential in improvement of cost and environmental characteristics are considered;

— the key design parameters of foreign HGTE and results of researches on the assessment of their efficiency by the advanced aircrafts using kerosene and LNG as fuel are presented;

— rational HGTE concepts are analyzed, numerical investigation of key design parameters of the mid-flight HGTE using different types of fuel, including gas and cryogenic fuels are performed.

2 most promising HGTE architectures based on the turbofan with high engine cycle parameters are defined in the paper.

— Architecture HGTE-1 — usage of additional supply of fan shaft by mechanical power from electromotor (EM). Electrical energy to supply EM is generated by external source, i.e. power unit (PU) based on solid oxide fuel cells (SOFC).

— Architecture HGTE-2 — usage of electrochemical generator (ECG) based on SOFC, operating in parallel with main combustor. Similar to HGTE-1 electricity, generated by ECG supplies EM, located on the fan shaft. Remaining heat generated by ECG supplies low pressure turbine.

Various levels of power ratio between the low pressure turbine and EM, providing a fan drive are considered for each HGTE scheme. Design parameters of effective schemes of HGTE are specified.

The forecast about prospects of development of considered concepts for 2030 and 2045 is made in the final part of the work.

In modern conditions of engine manufacturing the high demands for quality, reliability and durability are produced to the accessory gearboxes that provide the basic systems of the engine and the aircraft. The design of the accessory gearboxes should ensure its efficiency over the entire range of engine operation, from startup to maximum mode during the entire flight, including the possible evolution of the airplane. In this regard, the design of gears in addition to providing strength characteristics of the gears is also necessary to eliminate the possibility of resonant oscillations in the operating range.

If for some reason to get rid of the resonant modes is not possible, some special damping devices are used, limiting the resonance amplitude and stress level in the gear due to the work of the friction forces and the dissipation of energy fluctuations.

Experience in the use of dampers in mechanical engineering is quite limited due to the complexity of the analytical evaluation of their effectiveness by traditional methods and difficulties in manufacturing.

After the widespread introduction of numerical methods in the design of the gears in the 90 years of the 20th century the process of detuning was conducted much more accurate and efficient, and the direction of solution of resonance damping problems was weakly developed and now emerging techniques are in the making.

Therefore, in response to emerging operational requirements in FSUE «SMC of GTE «Salut» on the basis of the finite element method (FEM) the method of estimation of efficiency of the damper in the conditions of occurrence of resonant gear oscillation was developed and tested.

The main idea of the work is the realization of the possibility of the direct simulation of resonance oscillations of «gear-damper» system taking into account the friction on the surface of the contact interaction. The paper discusses the resonance vibrations of bevel gear with circular teeth for mode with two nodal diameters.

To achieve this goal, it was decided several intermediate objectives:

1. Creation a physically correct model of dynamic loading gear with solution of the contact problem in the gearing (the definition of the contact patch change);

2. Creation a contact model of the «gear-damper» system, which is optimized in terms of accuracy and dimension to perform labor-intensive dynamic calculations;

3. Determination of the new value of the resonant frequency of the «gear-damper» system on the mode shape under consideration by solving a dynamic problem with a special impact force causing the damped oscillations of the system in the interesting mode shape;

4. Carrying out dynamic analysis of the «gear- damper» system with excitement in its resonant oscillations, as well as the selection of the axial mounting gap on a damper that provides the necessary down force, in which there is no increase in the amplitude of resonant vibrations.

As a result of this work the new method was formed that allows to simulate the resonant vibrations of gears, to design and optimize various damping devices, preserving the amplitude of oscillation of the system to a safe level. Currently, the results of the implementation of the method was pilot-tested.

The work is devoted to the theoretical modeling of the behavior of the torsional actuator with a working body made of shape memory alloy (SMA). Devices of this type are very promising for the use in the elements of mechanization of the wing, empennage, aircraft or missile control systems.

In the frame of this work the problem of reverse martensitic transformation in the working body of the actuator with elastic counter-body, and fix condition for rods with solid circular cross-section made from SMA was solved. In both formulations of the problem with the heating rod made from SMA feels constrained deformation, the consequence of which is emergence of reactive stresses, which lead to an increase in torque.

The analysis is performed in the framework of the non-linear model of straining of SMA in phase and structure transition in coupled formulation with the possibility of occurrence of the structural transition. It is believed that stresses in this process are small enough, so that plastic deformation not occurs. The influence of the stiffness of the counter-body on the thermomechanical response of the actuator is studied.The dependences of the dimensionless twist and torque from the dimensionless parameter of temperature, as well as diagrams of dimensionless stress for different stages of the heating rod are obtained.

In this work it is established that in the process of heating the SMA actuator the bigger part of the cross-section rod is experiencing an increase in stresses. This requires to take into account the structural transition in solving problems of Reverse Phase Transformation (RPT) in the case of constrained deformation. The increase in the torque is a function of the stiffness of the counter-body and the torque, which active in case of Direct Phase transformation, and then the higher the torque then the greater the torsional stiffness has a counter-body. In the case of the RPT with fixed total deformations the phase deformations have transformed into an elastic deformation. This causes an increase in stress and torque. The maximum level of stress and torque at the end of the RPT is higher than the similar values for RPT with counter-body.

Implementation of gradiometry techniques for magnetic fields registration in the process of objects surface scanning is an effective approach to magnetic testing selectivity improving. We suggest vibratory inductive method based on induction transducer which sensitivity depends on the rate of change magnetic flux, coupled with turns of its inductance coil.

When vibrating coil moves along the surface under testing the first sign of blemish presence will be indicated by the first harmonic amplitude surge in the spectrum of induced electromotive force signal. With further movement of the coil towards the defect center a signal, representing a superposition of harmonics with frequencies ω₀ and 2ω₀, appears, with the tendency of increase in second harmonic amplitude and decrease in that of the first one. This is the second sign of a blemish presence. When the coil central section coincides with the center of fault stray magnetic field, the induced signal contains only second harmonic, which amplitude assumes the highest value. This is the third and the main sign of a blemish presence. It has a selective value with locality of flaw location.

In case of normal component testing circuit, the sign of blemish presence is revealed by a surge in the first harmonic amplitude over the fault surface.

The presented experimental data has been obtained through defected samples testing with a research facility, providing test sample movement against induction transducer according to three coordinates with resolution of 5 microns, as well as relative change of the transducer and test sample angular location.

Regulated rectifying units are the key functional units of power supply systems for different types of modern aircrafts.

At present thyristor and transistor controlled rectifiers are used on board an aircraft. However it should be noted that magnetic switches (MSs) preceded implementation of semiconductor switches used for electric power regulation and stabilization.

Highly efficient soft magnetic alloys development allows reconsideration of MSs energy and size efficiency indexes as well as their functional area spectrum.

The relevance of such class of devices studying is clearly indicated by the facts of their mentioning in foreign scientific and technical publications.

The foreign sources discuss design and implementation problems of such devices for modern and competitive aircraft power supply systems.

Special emphasis is made on application of advanced soft magnetic amorphous and nanocrystalline alloys for cores of MSs, which leads to a significant improvement of their electromagnetic properties.

New converter unit based on such magnetic switches is able to preserve the best features of its traditional prototype predecessor, in comparison with thyristor and transistor competitors, such as:

ease of implementation;

higher reliability;

better electromagnetic compatibility;

resistibility to special factors effect.

The authors have proposed new structures of aviation magnetic-regulated rectifying units (MRRUs), based on controllable single-winding saturable inductors based on modern amorphous (or nanocrystalline) alloys cores with rectangular hysteresis loop.

At present, there are no publications concerning engineering philosophy and design methods for this class of devices. Thus, the authors consider it expedient to offer their own methodology of design and estimation of controllable single-winding saturable inductors for proposed new MRRU structures.

The proposed method is universal and broad-based for transformer and transformer-less single-phase and multi-phase rectifiers.

We have investigated the levels of specific batteries (AKB) characteristics, which assure both the technical and economic feasibility of a cargo aircraft with electric propulsion system development.

The calculation of the characteristics of electric cargo aircraft was based on the following assumptions:

• The relative weight of the airframe and aircraft systems excluding the engine and propulsion system’s weight, the level of flight aerodynamic quality, and the required take-off power-to-weight ratio, in the first approach, do not depend on the type of the power plant;

• The number of engines and the type of the main propulsion, of the fan of electric aircraft, are the same as of analogues with the traditional power plant.

The dependences of the aircraft takeoff weight from the specific energy density of the battery and the relative masses of the onboard battery for a number of payloads and ranges are shown in tables 1.2.

Table 1. The cost of charging the onboard battery to deliver 20 tons of cargo to a range of 5,500 km electric aircraft

Table 2. The cost of charging the onboard battery to deliver 100 tons of cargo to a range of 9000 miles electric aircraft

The development of the mainline aircraft powered with electric power plant based on high-capacity batteries is technically feasible with the level of specific energy of on-board battery at about 4.5 — 5 kWh/kg (the maximum takeoff weight of the aircraft will exceed the takeoff weight of an analogue powered with conventional power plant by not more than 1.5 times, and aircraft performance will be identical). Upon achieving the energy density of on-board batteries over 6-6.5 kW·h/kg, the takeoff weight of the aircraft with electric propulsion system becomes lower than the takeoff weight of the aircraft fueled with hydrocarbon fuels.

Economic efficiency of applying the electric propulsion system in transport aviation primarily depends on the cost of the battery and its service life. By achieving cost-efficient service life utilization of ultra-battery at a level of 1.5-3 times lower than modern lithium-ion batteries demonstrate, the use of all-electric cargo aircraft will be economically feasible.

In the formation of innovative economy the scientific-technical progress, information technology, wide dissemination and use of innovations become the main factors of socio-economic development. Innovation is an integral feature of the modern economy. Innovation involves a large degree of risk and lack of complete assurance of success. Thus, particular importance is the study of projects of innovation, which depends on the reasonableness of the expert evaluations.

Under current conditions the research question is the impact of expert assessment on the decision of investors to finance innovative projects in the aviation industry. This article defines the essence and technology of expert assessments of innovation risk, and also considers the problem of the influence of expert assessments for decision-making on the financing of innovation project of the aviation enterprise. It includes an example of a risk assessment of innovative technologies for linear friction welding at JSC «Ufa engine industrial Association». It simulates an option of getting the opposite results. It is proved that expert assessment of innovative project is subjective. The main conclusions are that the important stage is the selection of the expert group, so as to obtain the most equitable assessments the maximum possible number of experts related to the sector should be used. They should represent different professional fields (engineers, of self-esteem experts is too subjective, and the results may be significantly overstated. It makes sense to use the evaluation of the professional level of experts by engaging outside experts, or by means of tests.

Initial information about an innovative project should be accurate and reliable at the time of assessment.

Taking into account all possible risks requires the great awareness of external and internal factors influencing the innovation process.

Development of small aircraft in Russia is possible with performing a wide range of tasks in interests of national economy, namely: development and production of inexpensive, highly effective, multipurpose light planes of passenger capacity categories, and large-scale production and a mass market launch of inexpensive, highly effective, multipurpose small aircraft with performances, operation cost, quality, safety and service guarantees for the purpose to replace the park of the outdated aircraft equipment and increase the level of small aircraft development.

The analysis of the development main directions of the small aircraft manufacture has to be directed on the solution to the following tasks: modernization of the existing park of the aircrafts (A) of small aircraft; mass production of planes of small aircraft; restoration and development of competences of the Russian aviation industry in the field of development of easy A; development of system of ensuring financial, technical and technological, personnel safety of the A manufacturing enterprises of small aircraft; organization of effective system of economic security of support for the production of Russian small aircraft; stimulation of demand for services of the airlines operating in the market of local air transportation.

The integrated approach to increase the level of economic security of small aircraft production has to include the analysis of the following directions: analysis of production specific; creation of criteria, indicators and their threshold values, and also the mechanism of providing economic security within state regulation of activity of the manufacturing enterprises of the aircraft equipment to formulate the organizational and methodical bases for formation of functional components of economic security of the enterprises.

For an assessment of economic security of the small aircraft manufacturing enterprises it is necessary to construct the model, which will provide the possibility for determination of economic security criterion value for the production of small aircraft. For creation of such model it is necessary to construct system of the indicators describing the analyzed object, considering not only the internal, but also external factors, for example, the structure and nature of changes in a domestic and foreign market, instability of an economic and political situation and others.

For an assessment of economic security condition of small aircraft production it is offered to use a complex of actions for prevention of threats within the offered methodical approach:

— it is necessary to make definition of system of indicators of economic security and their group on the main functional components;

— on the basis of the analysis and statistical data with the help of modern economic-mathematical methods it is necessary to estimate expected value of criterion of economic security of production of VS of small aircraft;

— to develop a complex of actions for prevention of the possible predicted threats;

— to make a complex assessment of current state of economic security of small aircraft production, which will allow to estimate quality of the developed package of measures and to consider possible mistakes and miscalculations for further planning.

Quality of the organizational and economic mechanism is characterized by functioning efficiency of the integrated structures of the space-rocket industry, consisting in ensuring life cycle of the created missile and space equipment. In the existing realities the need for improvement of the organizational and economic mechanism which methodological basis will be made by the principles of functioning of the integrated structures of the space-rocket industry is actual. Formation of a complex system of the principles is possibly at the accounting of a current state and tendencies of development of the space-rocket industry, features of activity of the integrated structures and qualifying standards to organizational economical mechanism of functioning. In the present work the three types of the principles that gives the chance to establish «trinity» of the principles of functioning of the integrated structures of the space-rocket industry are allocated. The first type — the principles of functioning of the integrated structures as hi-tech organizations: efficiency of functioning, continuity of the basic principles of the organization of production, subject and technological uniformity of the made production, compliance of the structural organization of corporation to object of production, a multi-nomenclature at single and small-scale type of production, stability of functioning of the enterprises entering association, unity of all stages of life cycle of the missile and space equipment. The second type — the principles of the structural organization of branch corporations: hierarchy, openness, stability, adaptability, interaction, management, proportionality, balance. The third type — the all-system principles: integrity, purpose, variety, complexity.

The established «trinity» of the principles could make a methodical basis for reorganization and improvement of the organizational and economic mechanism of functioning of the integrated structures of the space- rocket industry.

In this work a membrane wing of large aspect ratio in subsonic flow is considered. The wing consists of the front thin-walled beam and the rear beam connected in some transverse sections by ribs and by presstressed plane membranes located between the ribs. To calculate the deformed shape and aerodynamic characteristics of the wing we consider the plane problem for the wing airfoil.

The aerodynamic pressure acting on the deformed airfoil in subsonic flow is determined by using the exact solution in the series. The finite element method (FEM) is used to determine the small membrane deflections. The coupled problem of airfoil aeroelasticity is described by the system of linear algebraic equations for the node displacement of the FE-model. Solving this system we find the distribution of the aerodynamic pressure and the aerodynamic coefficients for the airfoil lift and moment. Besides these we find the boundary of the membrane divergence when the membrane turns up.

The example of the membrane airfoil is considered for which the numerical results are obtained in terms of the nondimentional parameter representing the ratio of the kinetic energy head to the membrane tension. Regulating the membrane tension by changing the rib lengths makes possible to control the aerodynamic characteristics of the airfoil and the wing on the whole without moveable aerodynamic control surfaces.

The results of this research can be used in design of the superlight high altitude unmanned aeroplanes with large duration of the flight.

The proposed method of the solution to a new aeroelasticity problem is original and the obtained results coincide with the exact solution in the particular case of the rigid airfoil.

The presented research is focused on the mathematical models of instruments, which are used to measure the altitude and speed parameters of modern aircraft.

The paper aims to investigate the possibility of increasing the accuracy of altitude and speed measurement during the highly maneuverable aircraft flight while solving the problem of perception of the total incoming air pressure.

The presented research is relevant due to the constant increase of the requirements for the accuracy of the altitude and speed parameter measurement during the flight. This increase arises from the necessity of ensuring the high level of flight safety and effective fulfillment of combat aircraft system missions. The matter point of the problem of accurate in-flight altitude and speed measurement consists in the following: various factors influence the initial parameters (pitot and static pressure, angle of attack and sideslip), which are measured directly with the appropriate on-board aircraft sensors. Multiple factors contribute to the total error of the primary airflow data perception. Among these the aerodynamic measurement error contributes the most. Therefore, the task of aerodynamic error reduction is relevant. The paper contains the solution of the abovementioned task for the pitot pressure instrumentation channel.

The SolidWorks and ANSYS software systems were used for simulation modeling. As a result the accuracy characteristics of the pitot pressure instrumentation channels of various pitot-static tubes were obtained. The characteristics were obtained at airflow Mach values ranging from 0,2 up to 2,2 for the angles of attack from 0 up to 30 degrees. The simulation was conducted for a standard pitot-static tube and several examples of various developed configurations. The paper adduces the results of the comparative analysis of standard and developed pitot-static tubes. After that the design modifications of the nose part of the developed pitot-static tube are substantiated.

The obtained results allow the authors to recommend conducting further research and tests of the proposed pitot-static tube model, which is equipped with a pitot pressure instrumentation channel that provides enhanced measurement accuracy. The implementation of the proposed design modifications of the on-board pitot- static tube would allow to determine the aircraft altitude and speed parameters with high accuracy for a wide range of speed and angle of attack values. This ability is especially important for highly maneuverable aircraft.

The development of the effective aviation armament system is one of the important issues that arise during the design of new and modernization of the existing military aircraft. In addition to weapon control system, sighting systems, etc. the aviation armament system includes the weapon launchers with the munitions that are installed on them. A multitude of technical decisions, which are aimed at the safe separation of the munitions from the delivery aircraft, is implemented to provide for the usage of aviation weapons. In general the safe separation refers to fulfilling the following requirements during the store flight in the vicinity of the delivery aircraft after its release from the launcher:

Technical decisions of integrated nature are required to provide for the munitions release. Such decisions define the essential parameters of all objects that make up the munitions separation system, namely, the «delivery aircraft — launcher — munition» system. These parameters include the following [2]:

Violation of the boundary values of the above- mentioned groups of parameters can result in aircraft or store damage, which may cause an accident or store mission failure. The separated store may become unstable, which can lead to mission failure or collision with other stores (during the separation of several stores) or adjacent airplanes of the operations group.

Technical decisions, which are aimed at providing for the usage of aviation weapons, are checked with the help of simulation of the store launch and subsequent analysis of store spatial movement in the vicinity of the delivery airplane. At that an important role is played by the reliability of interferential aerodynamic characteristics of the store-airplane system in the vicinity of the delivery airplane and characteristics of the isolated separated store.

Interferential characteristics in transport position and in the aircraft vicinity generally depend on a wide range of parameters such as:

The dependence of the interferential coefficients can be formally represented by the following function:

Thus the problem of determination of the interferential characteristics of the cargo-airplane system is multifactorial and includes all of the following: the design parameters, the location of the load on the airplane, the environmental parameters, the altitude and speed of airplane flight.

The described method allows the designers to determine the interferential aerodynamic characteristics of the delivery airplanes and stores of various shapes at early development stages without conducting the wind tunnel experiments. This can reduce the cost of wind tunnel tests and time, which is required for the optimization of the store aerodynamic configuration and its placement on the airplane, significantly.

The results, which were obtained by using the described method, can be used to calculate the parameters of separation of loads from the aircraft and separation safety evaluation.

The data, which was processed by using multidimensional (multivariate) interpolation, allows the researcher to analyze the interpolated dependencies based on all the factors that influence the interferential coefficient values. 

Lately a lot of popularity gained unmanned aircraft vehicles (UAV) for research current situation on the ground from the air. Simultaneously, the increased requirements for the quality of the image obtained with a camera mounted on the UAV. It is necessary to ensure offset images are not more than 1/3 pixel. As a consequence to the fore extends the problem of vibration damping obtained from the engines of the aircraft. To ensure high quality video it is necessary to extinguish the three-dimensional angular camera shake amplitude of units of angular minutes and tens of hertz frequency. Currently the most widely used two basic ways to solve this problem, applicable to aircraft: gyro stabilization and electromechanical systems. Options considered to have a not satisfactory characteristic of weight and size, or the high cost.

The purpose of this work is to construct a three-dimensional stabilization system that provides vibration damping of aircraft body.

Proposed a new approach to the construction of the stabilization system of the camera. This system is constructed on the basis of the mechanism of parallel kinematics (hexapod) that allows to get high precision system at its small sizes. Developed a mathematical model of a 3-dimensional stabilization system based on hexapod and solved the inverse problem of kinematics.

Simulation in Matlab allowed to obtain requirements for the hexapod actuators. Analysis of different types of drives has shown that necessary characteristics have actuators built on piezo actuators. Their distinctive feature is the low weight, great efforts on the output, high speed of working off control signals.

In this work was developed a mathematical model of piezo actuators that takes into account the following parameters: the total weight of the reduced mobility, the coefficient of elasticity coefficient of the inverse piezoelectric effect, the ratio of direct piezoelectric effect, the damping coefficient, the electric capacity, the internal resistance of the source of emf. Mathematical modeling shown the possibility of using the piezo actuators in 3-dimensional stabilization system.

Modern projects of Earth remote sensing systems (ERSS) provide combination of their target functionality for solution to a range of special tasks. That is why the spacecraft (SC) is designed as multi-purpose one and ERSS is also considered as multi-purpose. Therefore the problem of planning SC multi-purpose orbital constellation target functioning is multi-criterion. So, it is required to carry out at the same time solving of ERSS target functioning efficiency for several various criteria (or for one of the most preferable from them). The solving of such complex optimization problem is possible only during the development and application of specially designed methods for complex multi-criteria optimization adapted for multi-purpose ERSS functioning.

The main local and integrated indicators of ERSS target functioning efficiency such as operational efficiency of shooting information broadcast from SC on-board to ground stations, average daily volume of the shooting information broadcasted from SC to ground stations, frequency of Earth equator observation, the cumulative value of shooting information are considered in the article.

The general optimization problem statement for multi-criteria optimization of ERSS target functioning efficiency on the basis of operative planning is given in the article with the using of mathematical models and similar optimization tasks formalization presented in earlier authors publications.

The examples of solution to analysis and optimization model tasks operative planning of target functioning for ERSS constellations including from 1 to 5 SC with the project parameters «Resource», «Kanopus», «Meteor» are presented in the article. The results of the comparative analysis of target functioning of these constellations are presented as well. The results of multi-criteria optimization are given in ratio on different coefficients of importance of integrated indicators of ERSS target functioning efficiency. The corresponding recommendations about a choice of the ERSS constellations are formulated.

As screens of SVHI it is used the polyethylene terephthalate or polyimide film metallized with aluminum on one or both sides by the film thickness of 5-20 microns, the aluminum coating layer is 0.1 micron. Outer and inner surfaces of SVHI are sheathed by combined materials «Nyika-KPMA» or «Nyika- PAM-2» with a thickness of 150 microns.

To ensure evacuation and protection against electrostatic charges the films of SVHI are perforated by holes 2-5 mm with perforation pitch from 10 × 10 to 50 × 50 mm.

Laser perforation of SVHI films in comparison with other methods has a number of advantages: significantly increased productivity, increased accuracy of perforation. Presented method allows to quickly adjust the diameter and pitch of holes.

Jobs created by the laser system is based on the application of the principle of linear-angular scanning of the laser radiation, in which the movement is carried out by the laser beam on the surface of the material being processed. The principle of the system is realized through the use of linear angular scans, which consists of:

Laser perforation system can operate in one of two modes of perforation:

Calculations of the motion path of the laser beam in the scanning system and the estimation of the effect of errors on the accuracy of the process of perforation at various modes of perforation are carried out.

Comparing the hole-cutting mode it can be concluded that the start-stop mode provides a greater accuracy than by cutting apertures on a continuously moving film.

In the carried-out and planned missions of the spacecrafts (SС) to the Near-Earth asteroids (NEA) and other small space objects the laser devices are often applied. However, the range of tasks solved by them is usually limited by distance measurement at small ranges and precise height measurements during landing to NEA.

At the same time, the use of multipurpose laser equipment onboard SC allow to avoid failures and violations of processes of approach and landing to an asteroid (like it was, for example, during the landing of SC «Hayabusa» to asteroid «Itokava»), obtain various data on structure of a surface and substances of asteroid.

Here is a preposition on laser target equipment based on multi-functional 3D-lidar intended to be used on board of satellite missions to near-earth asteroids and other small space objects. 3D-lidar possible functions are as follows: precision range and velocity measurements, three-dimensional monitoring of asteroid shape and surface structure, precise descend of a lander, 3Dmonitoring of asteroid impact results, laser spectroscopy of substances of a space object, etc.

The lidar makes continuous measurements of the current range and radial velocity at all distances between SC and NEA. At small (some kilometers and closer) distances the lidar allows to receive detailed spatial «portrait» of object due to application of a threedimensional photosensitive matrix.

Process of satellite approach with NEA becomes complicated due to large (tens or hundreds kilometers and more) errors of preliminary data of NEA orbit parameters received from ground-based observation facilities.

In case of use of onboard SC of the «rough» distance-measuring instrument (for example, an optical direction finder or TV-camera) there will be a big uncertainty of knowledge of proper distance to NEA. It can lead to excessive costs of propulsion system fuel for operations of approach and landing to NEA.

Calculations of necessary liquid rocket fuel expenses of the satellite propulsion system during approach to NEA have been carried out. It is been shown that using an onboard laser rangefinder would allow to reduce fuel expenses approximately twice in comparison with use of only passive onboard observation systems.

It is also shown that during tests of deflection means of NEA from a trajectory dangerous to Earth the lidar can measure very small (up to 1...5 mm/sec) additional increments of the speed received by asteroid in the result of collision with the special SC-impactor.

Use of a multipurpose lidar may lead to some increase of satellite overall mass; nevertheless, it is quite compensated by increase of amount of the tasks solved by the onboard equipment and accuracy of their performance.

Analysis of the aerospace engineering needs and other modern self-contained technical facilities has shown that the aluminium-air (AA) mechanically rechargeable chemical current sources (CCS) can be effectively used as the main energy sources or emergency power systems.

This paper helps to solve an actual issue as the developing of the functional dimension-type series of AA CCS. All power plants gathered into one series via their characteristics (so called parametric series) are developed fr om one point of view.

The modular parametric range assumes that all power plants are built by means of consecutive or parallel commutation of the basic modules.

As the optimization parameter for the modular parametric range of AA CCS the electrode sizes were accepted, namely the size of the gas-diffusion oxygen cathode. The cathode usually has high over potential and its operation time lim it the operation time of the AA CCS due to its structure.

The evaluations of the different parameters of the modular range were also carried out in this paper. It is also shown that in the cases of difficult power consumption usually characterized by the cyclic higher power peaks the hybrid scheme of the power plant can be effective. Such scheme includes AA CCS and buffer accumulator battery (AB), for instance nickel metal hydride or lithium-ion batteries. AA CCS provides a basic power rate and recharges the AB, but AB covers the power peaks if it is necessary.

The evaluation of the required basic modules numbers has done and it is shown that the main factor which limiting the power rate of the basic modules is maximum achievable level of the one cell in the AA battery. This is determined by the technological limits of the large-area gas-diffusion cathodes production and the possibility to provide the proper conditions of the thermo- and mass exchanges.

This paper shows that based on the chosen optimal parameters it is possible to make 2-3 modulardimension-type series based on the 3-5 basic modules in each series. Besides, the basic module can include 1-6 or 1-8 battery depending on its power rate.

Issues related to measurement of particle size distribution of two-phase flow in the mixing devices of the combustion chambers of the air-jet engine by laser-optical methods are considered. Special attention is paid to the solution of problems that arise when a large distance to the measuring volume and polydispersed composition of the liquid phase. This is due to the fact that when working with real fuels, such as kerosene, it is necessary measuring equipment to remove possible distance, and the droplet size of the fuel in the combustion chambers, as a rule, vary widely. The block diagram and design of a laser measurement system are considered.

In the operation of the ring photodetectors, which are typically used in laser measuring sizes of particles, is replaced by a matrix of a digital camera. This allows you to create a system where the laser and reception optics do not have a rigid connection which is important in conditions of full-scale experiment. In this case center of diffraction pattern of each frame is set programmatically. The multiple times photographic registration of two-phase flow following by a specially designed procedure «stitching» frames is used to increase the dynamic range of the measurements . Also to improve the measurement accuracy characteristic curve of the matrix a digital camera is based, by which means is determined by the intensity registered by the matrix of light. It is shown that the using installed in the focal plane of the Fourier lens the additional light-diffusing screen allows you to extend the range of the measured particle size and to simplify the optical system of registration. In this case telephoto Fourier lens of large diameter were used to work at greater distances (in our experiments we used a lens with a focal length of 500 mm light diameter of 195 mm). A light-diffusing screen was made from sandblasted on one side of the glass by thickness of 3 mm and size 300 mm.

The block diagram and design system, laser measurement of dispersion of two-phase flows, which was used in full-scale tests are considered To improve the reliability of the measurements and to expend a measuring range, the emitter unit contains two laser emitters — red (wavelength 650 nm, type KLM-650/20) and green (532 nm, KLM 532-30-5). Depending on the set before the experimenter tasks possibly using different algorithms for processing the received scattering indicatrixes.

To assess the validity of the results obtained by using the system of laser measurements of particle size, comparative experiments with analyzer of particle size the firm Malvern were carried out The distance to the test object of research the analyzer Malvern was 4 times less. The comparison showed that the differences do not exceed 8%.

Also the results obtained in the experimental study of the mixer on a working body air-water with the formation of gas-droplet stream are presented. It was found that the investigated mixer with increasing gas flow rate at constant liquid flow rate Sauter mean diameter of the droplets is reduced from 140 to 90 ?m. At a constant concentration of the liquid phase and the increase of flow rate from 1.5 to 2.25 kg/s, there is a minimum in diameter which is in the area of consumption of 2 kg/s.

One of the major complex problems facing modern aircraft manufacturers is to increase energy efficiency, fuel efficiency and environmental performance of aircraft (PA), with an increase in their performance and expansion capability solved PA. To partially solve this problem in PA used and planned to use contactless bearing supports (CBC) that increase the speed of equipment aircraft, its environmental friendliness and reliability, while reducing the weight and dimensions.

From the analysis of works of Russian and foreign authors can conclude that in the elements and nodes of the aircraft used several types of CBC:

Table shows a comparison of the considered types of criterial CBC considering the application (aviation industry).

Table shows that the most promising type of non- contact bearings, to provide a reliable and efficient operation of the equipment of the aircraft, as well as provide the ability to integrate non-contact bearings in gas turbine aircraft engine is a hybrid magnetic bearings. In order to increase the effectiveness of their use in aircraft is necessary to solve a number of problems, such as an increase in the operating temperature of high-coercivity permanent magnets, the creation is not sensitive to the temperature of sensorless control methods rotor position, it is necessary to expand research on the development of high-temperature insulation materials, as well as to produce a study on control system synthesis capabilities hybrid magnetic bearings, and aircraft systems.

Comparative analysisof gas bearings and hybrid magnetic bearings

In addition, there are a number of unsolved theoretical problems of research of hybrid magnetic bearings, for example, there is no uniform classification of their design concepts and designs, etc. In other words, for the successful implementation of programs more electric aircraft, it is necessary to extend the common theoretical base on hybrid magnetic bearings.

The level of operational and technical characteristics of propulsion systems in modern aircrafts is determined by a combination of functional, technological and economic indicators providing high quality of aerospace industry items. There are some features in use of rocket engines as part of carrier rockets and spacecrafts which stipulate increased demands to the reliability of all its structural elements and the engine as a whole.

One of the main indicators of rocket engines quality and reliability is a stability of energy parameters in high- speed centrifugal pumps of fuel liquid components supply system. In engine production it is controlled by the results of modeling hydrodynamic pumps testing and assessment of energy parameters dispersion (head and efficiency) in comparison with their values set by design documentation.

In general case deviation of functional parameter in complicated hydraulic system is determined by an error of every influencing element, values of influence coefficients and error spreading character.

For high-energy pumps any hydrodynamic parameter deviation is stipulated by influence of deviations in geometric dimensions of details in flow section, deviations of details mutual position because of error in pump assembling process, errors of measuring equipment during tests and maintenance of a given test regime.

Analyzing the results of pumps tests it is obvious that the task of energy parameters stabilization can be solved by increasing of manufacturing accuracy of hydraulic path details, decreasing of errors in pump assembling process, raising measuring accuracy and calculation of the controlled parameters in units tests.

Experimental studies of specific investment of constructive and technological factors into the energy parameters deviation were conducted by standardization method when the pointed value of controlled parameter was result of handling from 9 to 27 attempts that allowed significantly increase the experiment veracity.

Experimental and analytical researches, calculation of coefficients of constructive and technological factors influence are made for pumps with closed centrifugal wheels.

As a result of analytical and experimental researches it is developed a calculation method of head and efficiency deviations, considering the influence of constructive and technological factors into the energy parameters stability of centrifugal pumps. Solving of this task promotes to the creation of high quality rocket engines with set level of reliability.

The goal if this work is functionality enhancement and software and hardware development for liquid thruster firing trials software and hardware complex based on algorithmic and hardware development for automated information and measuring and controlling systems (IM&CS) with the possibility of fault conditions functional diagnostics during liquid thruster continuous running.

Designed automated firing trials IM&CS allows testing of the existing and emerging liquid thrusters operating on ecologically clean propellant components and uses the function testing methods of faulty states during liquid thruster continuous running to determine emerging failures.

The applied methods are based on complex systems mathematical modeling and simulation methods, theory of measuring and measuring transducers, function tests theory, experimental research as well as bench trial and full-scale tests methods.

The developed fault conditions function tests technique uses fault condition classes as well as secondary features selected as fault characteristics. Fault conditions of each class causing new static mode of the thruster are c encoded by selected secondary features. The developed algorithms realizing the developed technique together with fault diagnostics algorithm allow assigning lined-up faulty condition to a certain class or classes in order to determine a failed sensory element or elements. We realized these algorithms in the form of software «Liquid Thrusters Fault Conditions Diagnostic Module». To develop the software we use Delphi package of Embarcadero RAD Studio XE.

Using the developed automated IM&CS included into thruster firing trials software and hardware complex we carried out tests of such thrusters as DMT MAI 202 operating on kerosene and oxygen (g); as well as thrusters DMT MAI-200-1P DMT MAI-200-7P DMT MAI-500VPVK operating on high-concentration hydric dioxide (HCHD 98%) and kerosene. We also carried out first stage tests related to developing prospective liquid thrusters operating on clear propellant components. Implementation of suggested algorithms allows expansion of firing trials software and hardware thrusters in the pipeline and prospective ones.

We carried out more than 400 experiments to test the developed software. These experiments fully confirmed its functionality with respect to fault conditions function tests.

The adequacy of the developed techniques and algorithms applied to fault conditions diagnostic subsystem for liquid thrusters in continuous running was fully confirmed in the course of liquid DMT MAI 200 thruster firing trials when fault conditions occurred.

Application of the developed IM&CS with the possibility of functional diagnostics of the liquid thrusters faulty states in continuous running allowed to reduce the amount of fire tests by 9% for DMT MAI 200 thruster development and studying.

In this paper we report on engineering of the automated ground-basedsystem for the measurement and control of meteoparameters of the atmosphere. Improvement of the existing hardware complexes and invention of the new automated gages on modern element base is the most perspective direction in meteorological instrument making.

Applying the considered gages it is possible to receive the following characteristics of controlled meteoobject in real time: height and temperature of the lower bound of overcast, water content, water stock and some other characteristics of clouds; temperature, pressure, humidity of air and parameters of a wind in a ground layer of the atmosphere; optical thickness of the atmosphere; coefficients of blackness of clouds of various form; assessment of transparency of the atmosphere; optical density of translucent clouds; density of a ground aerosol and sizes of its particles.

Measuring systems are united spatially and functionally in the NALRK general complex. This complex includes the following measuring systems:

Use of the ground-based automated lidar-radiometric complexwithresults of measurement will be expedient in TsAO of Roshydromet, RCfor Earth Operative Monitoring of Federal Space Agency, MNPO «Spektr», GPU «Mosekomonitoring», IFA n.a. A.M. Obukhov of the Russian Academy of Sciences, GGO n.a A.I. Voeykov and in other organizations connected with research and forecasting of a meteorological situation, and also with environmental monitoring in any region.

This paper is devoted to onboard brushless superconducting synchronous generators for full-electric aircraft. The new constructive scheme of superconducting synchronous machine is offered that contains claw-shaped poles and permanent magnets. The advantages of suggested constructive scheme of the machine are specified. It is shown the permanent magnets in a claw-shaped rotor increase main magnetic flux in an active zone of the generator.

The purpose of work is to get characteristics of machine and recommendations for choice of constructive data of machines on the basis of the obtained characteristics for further projecting.

Open-circuit characteristics and on-load characteristics of the offered synchronous machine with claw-shaped poles and permanent magnets obtained settlement and by practical consideration and reflecting advantages of application of permanent magnets in the offered design are provided. The obtained open-circuit characteristics confirm expediency of application of permanent magnets in claw-shaped rotor. On the basis of the obtained family of open-circuit characteristics the recommendations for a choice additional air gap value in a magnetic circuit of the machine at its design are given. In family of open-circuit characteristics there is a point of intersection of all characteristics. If the point corresponding to a nominal operating mode of the machine is below a point of intersection of characteristics, it is necessary to choose the greatest value of an additional air gap as in this case the main part of the main magnetic flux will be created by permanent magnets. If the point of the nominal mode is above a point of intersection of characteristics, it is necessary to choose the minimum value of an additional air gap should be chosen minimum as in this case the main part of the main magnetic flux will be created by excitation winding.

The on-load characteristic of the offered machine has good overloading ability. It is shown that in this design of the generator the superconducting excitation winding has to be used as the regulator of output voltage.

The aim of the work is to determine the settings of hardware delays in navigation receivers and satellite GLONASS transmitters in order to compensate ionospheric errors. The technique of the receiver calibration in this work is a method that does not use additional equipment to determine the settings of hardware delay algorithmic methods, taking into account the precise knowledge of navigation receiver position. The calibration allows one to retrieve values which are the difference of frequency-timing amendments not taken into account in the ephemerids. The assumption of the linear nature of hardware delay signals depending on the frequency allows one to determine the difference between the values observed delays in the L1 and L2 bands at zero, as well as announce changes lag of one type, in the ranges of L1 and L2, respectively. It is clear that these values are inseparable with the physical care of the timeline of the receiver and so will not have any impact on the assessment of the coordinates in the navigation solution.

The proposed method of calibration was applied to double frequency navigation receiver Javad TR-GNSS G3T. The study has shown the technique is sensitive to the accuracy of the navigation receiver. Using the same unbiased weighted difference allows one to apply the navigation algorithm for obtaining unbiased estimates of coordinates GLONASS receiver.

The subject of this research is creation of an algorithm for piezoelectric ring cell coefficients calculation, piezoelectric motor static and dynamic characteristics study, setup of technique for piezoelectric percussive type motor with piezoelectric ring cell model coefficients.

In the course of mathematical model setting up, we used decomposition principle, and split the model into three subsystems:

The updated algorithm for piezoelectric ring cell model coefficients calculation, based the dependences stated in reference [1] allowed to simplify design and analysis of piezoelectric motor characteristics.

The paper presents the annular piezoelectric cell resonant frequency tuning technique. We plotted Bode diagram to specify resonant frequency of the annular piezoelectric cell, which allowed us to determine piezoelectric resonator resonance frequency.

To create piezoelectric motor model we took into account the effect of the pushers represented by plate springs and described by compressed spring model. The model of the pusher also accounts for plate spring stiffness factor proportional increase. It allows ignoring trifling alterations of pusher to rotor slope angle. Comparative analysis of obtained characteristics and rated values lead to a conclusion on adequacy of obtained mathematical model.

Application field of the results of this research relates to the solution of the problem of aircraft recognition by optical systems to provide:

The aim of the work consists in determination of aircraft type by recognition system through its 2D image on photo-integrated matrix of visible, ultraviolet or infrared range in real time.

The proposed system of algorithms differs from its earlier developed counterparts mainly by recognition provision (with probability greater than 0.9) of 8 up to 12 types of military aircraft through their 2D dynamic optical images received by standard optical-electronic equipment and computing facilities.

Operating principle of the recognition system bases on the following principle. It singularizes an aircraft against the background of firmament with subsequent formation of its outer contour. Then the system compares this received real contour of aircraft image with reference contours of aircraft digital models from a given class stored in long-term memory. Reference contours herewith should be obtained at all possible aspect angles of digital models. Among all reference contours, the system must sel ect the one with minimum deviation fr om the real contour according to given matching metric. The basic means for time acceleration of aircraft recognition consist of algorithms with preprocessing and proper organization of long-term memory. Application of the introduced algorithm of real and reference images matching by Nikodims metric makes it possible to implement recognition of aircraft type in real time.

Working test of recognition system prototype hardware implementation on physical aircraft models and videotapes of real aircraft flights has shown high probability of correct recognition (0,92-0.98) with minimum time delays (40-280 ms).

In addition we have a possibility to solve the problem of correct recognition with probability 0,92-0.98 in practical formulation for 8-12 aircraft types using the following standard hardware capabilities:

1. Optical system parameters: frame frequency 24 Hz, receiving matrix size 720×576 pixels.
2. Computing device parameters: clock frequency 3 GHz, RAM 4 Gb, and LTM 256 Mb.

At present, the problem of object detection and recognition on video has many existed approaches to its solution. Methods herewith implementing object attributes description combined with machine learning algorithms gained maximum popularity. Any one of object characteristics, such as shape, color, or texture, can play the role of an object attribute. Some attributes describe local characteristics while other can characterize an object globally. Depending on attributes in use, one can decide whether such detector suits to solve the accompanying problem of object tracking. Moreover, in the case of a human figure — its posture analysis.

One of the existing and potentially useful global object descriptors is rectangle cover, which was earlier used in [1] to solve the problem of automatic classification of moving humans. Rectangle cover represents the aggregate of all maximum empty rectangles, inscribed entirely into 2D figure. These rectangles were calculated by morphological opening operation. Correspondingly, to determine all maximum rectangles of fixed size, one should calculate an opening operation of the entire figure for each size of this kind. The entire process was extremely time consuming.

Therefore, this approach appeared to be non-applicable for a large number of objects in picture. It worked very slowly even with three objects. The other problem lies in the fact that with application of a rectangle, we lose rotation invariance. The abovementioned problems defined the goal of the research: modification of method [1] in order to increase the speed and overall quality of the human figures detector.

The algorithm for solving this problem consists of the following. Video contains moving objects and some of them are humans. During the first stage, we separate silhouettes of all moving objects based on background model building using mixture of Gaussians. This method provides a high density of pixels within the sel ected moving regions, as well as allows us to separate the shadows from the objects.

During the second stage, we use continuous morphological models (skeletons) [2]. It allows to build morphological thickness maps (covers) (Fig.) of the objects effectively not for rectangular but for disk structuring element [3]. The advantage of using disk is its invariance to translation and rotation. This is a key step in the proposed algorithm. Disk thickness map is stable descriptor, contrary to skeleton, shown in [3].

Based on the thickness map object feature vector is generated. It is a set of disks, and five normalized parameters characterize each disk: radius, center coordinates, area, a number of unique pixels that were not overlapped by other disks. Further, classification stage begins. We need to determine whether the presenting silhouette, described by feature vector, is human or not. To do it we use algorithm containing extremely randomized trees. Every available disk is recognized within every tree. Each disk relates to a certain class by a majority of votes. Then votes fr om all disks are collected to make the final decision whether this silhouette belongs to a human being or not.

Qualitative evaluations obtained allow to draw a conclusion that the proposed method successfully solves the problem of moving human silhouettes recognition. Application of disk thickness map leads to increase of quantity of correctly classified human silhouettes and reduces the number of false detections due to invariant properties of the disk thickness map.

The proposed approach is superior in terms of performance to previously known method based on rectangular covers (12 objects are process at 10 frames

Human disk thickness maps examples per second versus 3 objects at 5 frames per second). One can use this approach for video surveillance systems, installed in small placements with small human flow.

The principal points considered in the article are focused around various aspects examining cause and effect restrictions of relativistic and quantum-mechanical origin on the transmission speed and the treatment of discrete information. In the introduction different approaches to Einstein-Podolsky-Rosen paradox and its actual generalization like effects, described by Bells inequalities, and also certain aspects of the problem history are demonstrated. Main content is devoted to features of a signal propagation in the closed slot microwave transmission line, metallic components of which are in the superconducting state. In case of coapplication of the classical model of the field screening in a superconductor and magnetic flux quantization in a closed superconducting circuit certain ascendant imperfection of the «usual» fluxon theory, connected with some contradictions, is demonstrated. Correlation of quantum fluctuations and propagation feature of pulsed magnetic disturbances in a closed extended superconducting circuit are analyzed in the context of Einstein-Podolsky-Rosen paradox. From presented evaluations it follows that the process of signal reception at the end of the superconducting slotted line with conservation of quantum coherence is possible under corresponding selection of detector parameters, such as SQUID. On the whole the performed consideration shows us that the macroscopic quantum phenomena constitutes not less abundant list of effects, than traditionally reviewed schemes of EPR paradox realizations, based on microsystems in the form of photon pairs, which are born in the entangled states. In the conclusive part the possible schemes of an experimental study of postfactum influence effects are considered, when the effects of influence on interference in a single macroquantum system from classical positions may be interpreted as the posteriori sequence of consequence-cause. «The post factum action» is realized as an impulse disturbance of the optical path length, introduced into one of shoulders of Mach — Zander interferometer. For «experimental facilitation» of the task in both shoulders of set-up optic delay on the basis of photon crystals, which is providing time lag, corresponding to the light propagation per cents of kilometers, are introduced.

The problem of the competitiveness evaluation of aviation technology at the stage of making decision about buying and formation of the Park is relevant in connection with active development of transport system in russian regions and cities.

Methods of the estimation of competitiveness that currently proposed in the literature and other information sources are not designed to cater for the increasing needs of customers, the number of private consumers aircraft increases.

In the presented paper we describe the developed method in determination of the integral competitiveness of the helicopter in helicopter group of a particular class. The approbation of this method on the example of civil helicopters is given as well.

The integral market competitiveness of the helicopter is encouraged to understand as the result of the comparison of each model helicopter candidate for purchase with the sum of analog — helicopters of different manufacturers, are on the feasibility and consumer characteristics, taking into account the blur variations in consumer preferences.

It is proposed to use the method based on the theory of fuzzy sets, prioritizing in the assessment scales analyzed analogues helicopters for making decisions regarding the level of overall market competitiveness.

The result of the competitiveness evaluation in the article is invited to consider obtaining a competitive range of helicopter models, prioritized on integrated index of market competitiveness.

Purposes

1. Determine the relevance of the organization modernizing of repair manufacture at «Chernyshev MME, JSC».
2. Analyze the segmentation of market of aircraft turbine engines for the light fighters.
3. Analyse the perspectives of development of each engine modification that is produced and repaired at the companies.
4. Carry out the system analysis in organization of repair manufacture for the aircraft engines at «Chernyshev MME, JSC».
5. Find out the opportunities to create the mechanism of organization of repair manufacture for the aircraft engines within the indicated conditions of external environment development.
6. Determine the tasks of optimization of process routing for the repair manufacture.

Design/methodology/ approach

Using the methods of competitive analysis and segmentation of market to determine the perspectives of development of each engine modification that is produced and repaired at the enterprise. Conducting the research of process in organization of repair manufacture on example of «Chernyshev MME, JSC». Investigation is made by decomposing the object using the system approach and SADT-modelling. Using the method of process routing to determine the sequence of process operations and production equipment used. Using the methods of making the management decisions to determine the optimum load of equipment and route of transferring the product fr om one lim it to another.

Originality/value

The following statements and results presented in this work are new:

1. Perspectiveness of development and expediency of forming the mechanism in organization of repair manufacture for the aviation components.
2. Trends of engine building market development are identified.
3. Decomposition of research object with using SADT-modelling is made.
4. The unit of selecting the optimum process routing is identified to be the main component of the organizational mechanism of repair manufacture.

The paper proposes a method for ensuring the required accuracy during the aircraft units mounting. The method was developed for the cases when the dimensional coupling structure of the assembly is represented by multi-link coupled three-dimensional chains. In such cases the assembly accuracy is attained via the adjustment method. Any change of one of the varied dimensions (of the so-called «variable compensators») during the attainment of the accuracy of assembly of such systems leads to the changes of all resulting parameters of the assembly precision. This is due to the fact that such dimensional chains are coupled. The output dimensional parameters are usually measured indirectly due to the difficulties of direct measurement of the relative spatial positions of the large-sized and heavy units. The adjustment process for such systems is implemented according to the trial-and-error method. The sizes of each variable compensator are consecutively changed and all output accuracy parameters are measured at each step of this process. The spatial positions of the adjusted points and output dimensions control points are separated by a relatively long distance. Therefore the adjustment requires the joint work of the assembly crew. This in turn leads to a considerable increase of the manufacturing cost and duration of the mounting and assembly.

The presented method implies the development of mathematical (numerical-analytical vectorized) model of the mounted units. This model describes the whole geometrical structure and all dimensions of the assembly. The numerically stable algorithms, which allow the operator to solve both direct and inverse problems, are constructed for this model. The database with the virtual assemblies is created by using these algorithms. Random values are assigned to all internal dimensions of each assembly in the database. These values are distributed according to a certain law within the given tolerance ranges. Each record of this database contains information about the changes in the variable compensators, which are required to attain the required output dimensional precision of the assembly. During an actual assembly this information is taken from a similar virtual assembly and used to tune the output assembly precision.

The proposed method is illustrated on the example of engine mounting technology for a single-rotor helicopter with two engines. The paper concludes with a discussion of the advantages and problems of implementation of the technology, which is based on the offered approach, in small- and medium-scale production of aircraft and rotorcraft.

A gliding unmanned aerial vehicle (GUAV) is considered. This vehicle is equipped with a high-aspect ratio wing and pneumatic control actuators. These elements make the long-range gliding flight of the UAV possible. The GUAV is the newest type of aircraft with improved performance.

The problem of guidance onto moving targets is very topical. The paper considers the problem of guidance onto moving targets under the conditions of flying along the designated route. The route is defined by navigational guidepoints. The number of these points may vary. The paper suggests an algorithm for generation of the initial states sets for the GUAV with the purpose of fulfilling the given mission. This paper is the development of papers [4-7]. In contrast to [7] this paper suggests a modernized and improved algorithm for the generation of the initial states sets for the GUAV. This algorithm consists of several stages: the stage of determination of the possible states set for the target (the point where the target is situated); the stage of determination of the critical points within the previously defined target states set (the critical points are the points, which are hardest to attain for the GUAV under the given initial conditions); the stage of generation of the initial states set for the GUAV. In this case the generated initial states set would be created as the intersection of GUAV initial states sets for all of the chosen critical points.

The target states set can be geometrically represented as a circle for the case of constant target speed. This paper illustrates some special cases, for example, when the target speed is low. In this case the target states set would transform into a circle with small radius, which can be geometrically represented as a point.

The paper considers the problem of simulation and identification of the nonlinear controlled dynamical system on the example of angular motion of a highly agile airplane. The proposed semi-empirical dynamical neural network-based model combines theoretical knowledge about the simulated objects with the tools of model improvement, which perform training of artificial neural networks. The theoretical knowledge is initially presented in the form of differential equations, which describe aircraft angular motion. Afterwards this knowledge is transformed into modular dynamic neural network representation. Insufficient accuracy of the theoretical model is improved via identification of the (nonlinear) aerodynamic force and moment coefficients. These coefficients are represented by the MLP (multilayer perceptron) modules, which are embedded into the complete model. A representative training set is generated to achieve the acceptable generalization capability and reduce the time, which is required to collect the data. This set is generated via special procedures of control input signal optimization. Training of dynamical neural networks on long data sequences has proven to be very difficult. There are several reasons for these difficulties, which include the effects of vanishing and exploding gradients as well as presence of spurious the training of the dynamical neural network model is performed in incremental fashion as follows. In order for the gradient learning method to attain the deep enough minimum of the error surface the training must start with the parameter values, which are close to that minimum. The problem of obtaining such parameter values may be treated as a problem of searching for minimum of a similar error surface. Thus a sequence of error surfaces with the following properties is constructed: a) the first error surface is simple enough to search for minimum with any initial parameter values; b) each next error surface is similar to the previous one; c) the sequence converges to the original error surface. Thus the original training problem is decomposed into a sequence of more simple sub-problems. Computational experiments confirm the efficiency of the proposed approach: the modeling errors for the test set are

where ,  are the angles of attack and sideslip; p, q r are the angular velocities about X, Y, Z axes.

Aerodynamic coefficient estimation errors for the test set are

where are the coefficients of body-axis components of the aerodynamic forces and moments.

The paper studies the physico-mathematical models of the flows, which are based on a system of moment equations. The main shortcoming of the moment equations systems is their short-wave instability. This flaw limits the use of the models in hypersonic regions of the flow.

In [1] the system of Grads moment equations is extended to polyatomic gases. The publication offers a method of constructing such system with a smaller number of additional assumptions. The obtained 24- moment system has the short-wave instability properties, which are similar to those of Grads 20-moment system.

Publication [2] showed that the main cause of the short-wave instability consists in the fact that the expressions for closure moments (4-th order moments), which were obtained by using the approximating distribution functions, do not satisfy the closure moments equations. This means that there is a discrepancy between the local and balance expressions for the closure moments. Thus the publication proposed the methods of the short-wave instability reduction.

One of the methods was connected with the expansion of the moment equations system. Additional members (matching additives) were introduced into the local expressions for closure moments. The resulting expressions were ins erted in to the closure moments equation. The differential equations of matching additives were obtained as a result. The original 24-moment system together with the matching additive equations created a 45-moment system of equations. This system allowed the research team to expand the region of solutions that are free from short-wave instability.

This work considers the method of [2], which is connected with the selection of the members with the highest order of magnitude from the matching additive equations. The paper considers two techniques of selecting the members with the highest order of magnitude.

The first method consists in excluding the non- equilibrium stresses even from the members with the highest order of magnitude. The moments of the 3rd order, which are included into the members with the highest order of magnitude, are left unaltered.

The second method implies excluding the deviatoric part from the moments of the 3rd order in addition to the above-mentioned alterations.

The algebraic expressions for matching additives are obtained as a result of this exclusion. The obtained algebraic expressions are included into the differential equations of the 3rd-order moments.

The usage of the proposed methods is illustrated on the example of the solution of the problem of the shock wave profile determination. This example shows that the obtained flow model allows the researcher to expand the region of solutions, which are free from the short-wave instability, close to the hypersonic Mach numbers ( M≈ 5 ). The second method of selection of members with the highest order of magnitude has significant advantages.

In this paper the method for structural-parametric system construction of optimal facial features flying objects is considered. For the case of a homogeneous target set the structural parametric synthesis is based on a statistical approach. For the case of multi-dimensional, heterogeneous set of targets the structural parametric synthesis is organized in the form of global search of the objective function.

Multipurpose feature of flying objects is determined by the variety of tasks performed by the system, the conditions of its operation and by set of system elements. It causes a number of problems which are far from being resolved now. The main problem is related to the assessment of the aircraft effectiveness, when it is difficult to decide what rules should get an overall effectiveness assessment of the aircraft system performing predetermined range of tasks. Volume of processed information in the system is much more than information capabilities of its individual components and thus the system should use a set of elements. Construction of multi-purpose system is rational distribution of a given set of targets between the centers of X strategy for flying objects building system. It is shown than external replacement of X  by some «calculated response» is unacceptable, since «design characteristics» often lead to a significant systematic error, and in some cases they may not contain.

In this paper we present a solution to the model problem in which it is necessary to optimally distribute the targets between cruise missiles group.

Purpose

This article discusses software complex of flight planning of unmanned automatic spacecrafts. Complex was developed based on methodology of automatic flight planning [1] describing the procedure of automatic variant of flight plan making and its verification using simulation model of Onboard Control Complex (OCC) [5].

Design/methodology/approach

Discussing way of software complex design includes formulation of complex creation problem, development of requirements and formulation of solving problems, structure definition and procedure of structure elements interaction. The main tasks solving by software complex are:

1. Automatic flight plan creation;
2. Automatic flight plan checking;
3. Transformation of flight plan into different views using during the plan realization by space system elements including spacecraft and Ground Automatic Control Complex (GACC).

In accordance with the problems being solved the software complex has been divided into four parts.

The first part of complex solves problems of general and daily flight plans. They contain data on the spatiotemporal spacecraft state, on executing control operations, spacecraft hardware modes and facilities of GACC, and other information related to the management of spacecraft.

In the second part of the complex the checking for the presence of a flight plan errors and the possibility of implementing systems and devices spacecraft is performed. To do this, the complex includes a simulation model of OOC, which helps in modeling work of OOC of spacecraft when the flight plan is executing.

In the third part of the complex it is prepared the plans for private systems and devices of spacecraft and GACC facilities ensuring the implementation of the flight plan.

The fourth part of the planning complex ensures the organization and management of data generated as a result of the first three parts. To do this database and developed database administration software is used. The proposed structure of the software system proved its functionality and allowed to take into account all the features of the phase of drawing up a flight plan implemented over an extended period of time.

Findings

The task of creating software complex automated flight planning of unmanned spacecraft allowed to formulate requirements and complex tasks, build a structure that meets the characteristics of procedures for the creation of the flight plan. In the result of this work a software complex that allows to solve the problem of the automated task of plan creation and its automated verification has been developed.

Due to created complex it is possible to simplify the procedure for flight planning, reduce the number of mistakes made by the specialist in the preparation of the plan, to make it more visible and comprehensible. Created software complex was introduced into practice management and used in the Mission Control Center (MCC) to manage automatic spacecraft of scientific and socio-economic purpose.

Research limitations/implications

Proposed automated flight planning software complex can be applied in the management of automatic spacecraft of scientific and socio-economic purpose flights.

Originality/value

1. Modern development level of technology automatic spacecraft control shows that the «manual» method of forming a flight plan becomes ineffective, and in most cases — impossible without the use of automation.
2. Feature of flight preparation procedure is that the plan is formed in several stages successive approximation for 2-week, daily and circuit intervals. Therefore, when planning in automated way it is necessary to consider the possibility of plan correction at any stage of planning.
3. Developed software complex for flight planning of automatic spacecraft has a modular structure, it provides a solution to problems of the automated preparation of the plan and its validation using a simulation model of the OCC. It also provides visibility into the process of the plan forming and plan correction by the operator, complies with the requirements of universality, so it can be used for different spacecrafts.
4. Due to created software complex it is possible to reduce the complexity of the planning, to reduce execution time from 1.5 days to 40 minutes, to improve the quality and efficiency of flight plan creating, to reduce the number of errors related to the human factor. Thus, in contrast to the «manual» method of making a flight plan, which requires the involvement of several specialists management group for the automated preparation method it requires only one person.
5. Software complex of automatic spacecraft flight planning was introduced at MCC and was used to control the flight of automatic spacecrafts of scientific and socio-economic purpose.

The joint project ExoMars (ESA, Roscosmos) consists of two missions in 2016 and 2018 with the descent modules (DM) weighing 600 and 2000 kg respectively. Calculations for Mars entry optimal parameters of flight path modules were performed. The optimal nominal input parameters were obtained as velocity 5.817 km/sec and the entry angle 11.79 degrees. Scheme with an aerodynamic braking deceleration portion and a portion of the parachute braking subsequent embarkation with the soft landing engines was chosen. Calculations showed that the aerodynamic deceleration provides a decreasing in velocity of descent module more than in eleven times. DM velocity reaches about 500 m/sec at 220 and 200 sec (for mass of landing modules 2000 and 600 kg respectively). In this case Mach number is equal to 2.1 and 1.9 (respectively), which allows to use a supersonic parachute. There will be applied parachute of 12 meters in diameter for deceleration of descent module weighing 600 kg. For braking descent module weight 2000 kg it is proposed a two-stage parachute system with diameters 15 and 35 meters of canopy. Calculations showed that the landing deceleration parachutes provide decreasing in modules velocity to subsonic speeds of 50-60 m/sec and then begins the final stage of braking by a soft landing engines. Calculations showed that in condition of super-and hypersonic speeds Reynolds number varies in the range of (1-1.6)?106. Entry trajectory of Mars descent module trajectory in «Mach-Reynolds» coordinates shows the range of modeling parameters, which have to be provided at the stage of descent module experimental testing.

Gas actuators play a great role in modern aircraft building. However, a lot of difficulties occur while designing actuators with short operating time. The lack of adequate mathematical models results in the higher costs of experimental researches needed to optimize design.

A model of gas actuator dynamic, generated by considering gas mass and energy balance in actuator cavities is presented in the paper. The balance equations are derived by integrating general continuity and energy equations over the volumes of actuator cavities and by using Gauss divergence theorem.

The main equivocations of the model, which occur in actuators with short operating time periods, are revealed. The methods for eliminating these equivocations are given. The modeling results obtained, using adjusted model, are stated.

In this paper it is shown that in case of short operating time actuator a zero-dimensional model of a gas actuator yields to substantial errors. These errors could be corrected by 3D modeling and by using the experimental data.

The non-uniform pressure field inside the actuator is the main factor that influences onto actuators dynamic. The second factor is the velocity profile. And the third important factor is a width of technological gaps in an actuator since these gaps define mass flow rate being exhausted from the actuator.

Authors reveal that the first and the third factors are of utter importance, while the second one could be sometimes neglected at all.

In this paper authors use two extra coefficients derived from the results of CFD-modeling and test data to adjust the mathematical model of the actuator.

The model after modernisation could be used for design optimization of the actuator.

In this article it is shown that the accuracy of pyrheliometers installed in artificial Earth satellites and applied for measurement of solar constant or total solar irradiance (TSI) is insufficient for measurement of long- term variations of TSI at the level of 0, 1% per years. The calibration method of pyrheliometers by feeding on them monochromatic radiation of several lasers with stable energy characteristics and different wavelengths is considered. Three methods of accurate laser power measurement are debated: with the help of cryogenic vacuum radiometer, silicon planar photodiode with additional transparent electrode on illuminated area and method of laser PIN- photodiode comparison.

The relative error of a measurement of laser power by the first method is 0,01 — 0,03%, and it is estimated by the computational way. If one applies silicon planar photodiodes, it strives growths of quantum efficiency up to 100% (on the statement of authors) by the feed of growing potential on the transparent electrode. Some independent researchers consider that the relative error of a measurement of a laser power by the given method could make up 0, 3%. The third method is based on the application of planar photodiodes with external or embedded temperature detector.

If one feeds the calibrated electric power on such photoconvertor, then it is translated into heat flow copying as on coordinates as at whiles the heat flow induced by dissipation of radiation in photoconvertor due to a drift of charge carriers initiated by the radiation.

Because of this the error of laser power measurement drops to 0, 001%.

Two methods for control of residual connected with inequivalence of optical radiation and electric current conversion in heat flow are pointed out.

First is based on the use of two operating modes of photodiode — short circuit and idle pitch, second is realized by the variation of thermal environment characteristics near illuminated surface of photodiode.

То increase the right verification of TSI variations it is offered to create an imitator of TSI and to strive a suitable reproducibility of Earth measurements in conditions analogous to orbital ones. The principle possibility of application of TSI monitoring results for forecasting flares of solar activity is mentioned, which should increase the reliability of aerospace electronic systems.

The object of research is high-stressed parts of high-temperature gas-turbine engines (GTE) and gas-turbine plants (GTP) with the complex of branched channels.

The aim of the research is to develop a methodology for creating and designing of cooled structures in high-temperature gas turbines of GTE and GTP of 5-6^th generations by using the experimental database to ensure the competitiveness of aircraft engines due to innovative technologies used for their creation.

The complex program of thermal designing the heat- stressed parts of the gas-turbine engines and gas-turbine plants for increasing the reliability and lifetime of the engine is developed. The methodology of mining the heat-stressed parts on the stage of the design- technological adjustment of cooled parts is developed. New criteria relations dependence characterizing the thermal and hydraulic conditions of perspective systems of heat exchange intensification of the cooled gas-turbine blades are obtained. The data bank for channels with complex configuration of modern systems for intensifying heat exchange in the high-temperature gas turbines of gas-turbine engines and gas-turbine plants is compiled.

Scientific novelty of the research as follows: the complex program for thermal designing and improving the structures of heat-stressed parts in GTE and GTP aiming to increase the reliability and service life of the engine by using the highly informative methods of research and by controlling of their functional parameters is developed.

The developed bases of technology of working off of the constructions of heat-stressed parts on the stage of the design-technological operational development allow to reduce the complexity and expense for the creation of prototypes of the aircraft equipment.

The received samples of the design elements of the aircraft equipment can be used in the development and creation of modern and perspective gas-turbine engines, and also in the process of building of heat-engineering devices.

For calculating the indicatrix of infrared radiation (IR) signature propulsion aircraft developed. A new technique based on the Monte Carlo method for calculating the signature of infrared radiation produced by an aircraft propulsion system is developed. This method is used for solving problems by simulating random variables. Calculations are performed by example of AL-31F engine.

The turbulent gas flow inside the nozzle is simulated by solving Navier-Stokes equations. For calculating the heat transfer in the wall of the nozzle the grid and geometry of the nozzle are generated. The temperature distribution inside the aircraft engine result is obtained. During simulation the inner surface of the propulsion system is modeled cones, cylinders, hyperbole, turbine ring, the central body ellipsoid. Infrared radiation are varied in a wide range of angles. In order to reduce this level, we offered to implement a special screen, which is a cylindrical surface extending over the nozzle exit. The screen diameter is large than the diameter of the nozzle and can significantly reduce infrared radiation in the range of angles from 10 ° to 90 ° depending on its configuration. The results of calculations for the

screening element with a different radius r, the length h and the distance to the nozzle exit plane H is shown in figure.

But the usage of such screens is difficult in existing aircrafts. Its overall efficiency is increased slightly with screen size increasing (since the length of 0.85 m). By using a special screen design there is a significant (98%) reduction of infrared radiation in the radial direction. But the level of radiation along the axis is increased. In this screen geometry (length and diameter) are important parameters. The choice of screen geometry defines by designers of airplanes. For the terrestrial infrared equipment matters most radiation in the radial direction and using of such device may be effective.

Under a spacecraft (SC) with long-term lifetime we mean a vehicle with an active shelf operating time of about 10-15 years. SC should be considered as a unified energy system which consists of a number of interactive subsystems. The main part of the SC power system includes an energy unit (EU) which provides the energy for all power consumers. Only solar or nuclear energy can be considered as primary energy for the vehicles under examination. At the present time it is the basic primary energy source for the low-earth SC and for the Sun directed probes. The main way for developing of the space solar power usage is the direct photoelectric conversion of the solar radiation. The development of huge power energy production (more than megawatt) will require the application of the nuclear power sources.

The main part of the SC EU is an energy accumulation system based usually on the storage battery (SB) that presents the major part of the entire EU mass and often determines the battery life. The lifetime of the modern SB depends on the operation conditions and it ranges from several hundreds of charge-discharge cycles (Ag-Zn) to dozens of thousands of cycles (Ni-Cd, Ni- H₂). The required SB life time depends on the SC orbit and is achieved by the deducting the discharge level.

The life cycle of modern capacity storages (supercapacitor) is greater by an order than it is for any other SB but their specific energy characteristics are much lower. Due to the increase of the suitable SB mass in case of the discharge level reduction and quite big amount of cycles the capacity storage characteristics will make them a promising energy storage system.

At the present time the Li-ion SB is widely used in SC on-board power systems. The necessity to use Li- ion SB as a part of the SC is caused by the increasingly stringent requirements for the increase of the specific power loading and reliability of the on-board power source. At the present time the life time of the Li-ion accumulators is insufficient for the SC under examination (near 1-2 thousand of charge-discharge cycles). The possibilities of the life time increase by means of the discharge level reduction (by analogy with Ni-Cd SB) have been studied insufficiently. However the continuous improvement of these sources allows to hope that of the required life time characteristics in the future.

The solution of the task of propellant unloading fr om the tanks of launch vehicles is extremely important to solve the problem of the minimal propellant residuals determining.

This fuel residuals prevents of tank pessurization gas enters in the feed line of liquid-propellant rocket engine.

The pressurization gas gets into the feed line from the fuel tank through the drain hole at the stage of complete emptying of tanks and the «broken» of the free surface of the fuel due to the formation of the funnel. The critical level of fuel preventing gas bursts out into the feed line is largely determined by the drain hole structure and the Froude number. The formation of a gas-liquid mixture is occurred due to the collision of the gas flow from the tank with the surface of the fuel in the feed line. The entrapment of the pressurization gas and its penetration into the fuel are results of this collision. The propellant residuals in the tank and the feed line are determined by numerical simulation of this collision. The numerical simulation of the liquid detanking process is done with the help of ANSYS CFX software package.

To solve this task a combination of the method of calculating the current of continuous multiphase medium with free surfaces (Volume of Fluid) and the method of calculation of the current of multiphase polydisperse medium, wh ere a liquid phase is continuous and a gaseous one is represented by a set of bubbles of different sizes (Multiple Size Group model), is used.

Until the moment of penetration of the gas jet from the tank at the free surface of the fuel in the feed line, the calculation is carried out by the method of VOF. The calculation of a gas-liquid mixture in the feed line is carried out by the MUSIG method.

The results of the calculation are well correlated with the experimental data.

At present time the attack unmanned aerial vehicles (UAV) with large take-off weight is the most interest among the wide variety of UAV. These UAV can carry a certain amount of armament besides a reconnaissance equipment. In the USA the development of ship-based unmanned aircraft systems (X-47B, «Phantom Ray», «Sea Avenger» and «Sea Ghost») is leaded more actively in the UAV class. The European projects, such as «nEUROn» and «Taranis» and Chinese project («Lijian») are also known.

The development of the UAV with attack opportunities in our country began at the end of 1990 years. The development of aircraft of such class engaged «MiG» corporation («SKAT»), the design engineering bureau named Yakovlev («PRORYV») and others, but presently none of them reached the fase of serial production.

The requirements to the power plants of such aircrafts are formed on the basis its implementation and solwing tasks. This includs the regimes of takeoff and landing, acceleration, maneuvering in horizontal and vertical planes and other.

The analysis of influence of design parameters of both the engine and aircraft, such as bypass ratio, takeoff thrust, the gas burner exit temperature, the overall pressure ratio and the wing area on the course of UAV performances is given in this work.

The ultimate aim of mathematical modeling at the aircraft design and its power plant is the optimal values obtaining of its design parameters and optimal control schedules.

With third-order astatism a phase-locked-loop frequency control system can be described by the following equations

Here is the adjustment period; i are phase and frequency of NCO (Numeral Control Oscillator) after completing the i-th adjustment period;

is the number in the third integrator of the PLL loop filter equal (in a steady-state mode with no noise) to the derivative of the incoming signal frequency ; is the signal at the output of the PLL discriminator; are coefficients of the loop filter determining dynamic and noise properties of the system.

These equations are still valid for the second order if .

Incoming signal phase instability results in PLL tracking errors which we will name dynamic ones. Here is the phase of the incoming signal, is the NCO phase at any random time instant t.

The paper considers dynamic errors in a steady-state mode. Such mode occurs after transient processes completion when , where n is PLL astatism order, and m is the order of the polynomial describing . In steady-state mode the tracking error averaged over the adjustment period when and when . However, tracking error instantaneous magnitude  can vary within the adjustment period, i.e., intra-periodic fluctuations occur.

The paper presents analytical equations for such intra-periodic fluctuations.

The results obtained can be implemented for PLL systems design to determine ultimate dynamic effects possible for such systems.

A problem of on-board equipment electric power supply always arises during prospective aeronautical complexes design. Judicious and justified selection of electric power system contributes to substantial time saving during design and realization of new airship.

Validation of electric power system type selection for aeronautical complexes, based on automated aerostats or unmanned airships, is generally carried out by expert analysis main mass-energy characteristics of power sources. Basis of such comparison and selection of optimal electric power system is relied on the condition of maximum values of energy performances with minimum weight and costs.

Due to poser sources, small storage capacity present-day electric power systems cannot provide required flight duration and power loading of an aerostatic aircraft.

Moreover, increasing the number of power sources certainly leads to addition to flight weight and correspondingly to payload weight reduction. Prevailing conflict resolution is provided with prospective electric on-board power systems application, and in particular such systems that can maintain on-board equipment operation of an aerostatic aircraft, as well as operation of purpose-oriented payload. These systems should have, herewith, improved energy characteristics and light load.

We carried out analysis of main mass-energy criterions of power sources for aeronautical complexes electric power systems in comparison with home-produced chemical current sources, as well as prospective electrochemical generators and photovoltaic power systems.

The results of the analysis show:

1. Chemical current sources, namely batteries, do not meet the requirements imposed by prospective aeronautical complexes power systems, due to low energy characteristics and heavy load. In addition, for extended operation at altitudes higher than 16 kilometers chemical current sources need heat insulation by way of water heat storages. Implementation of heat insulation leads to 30% increase in electric power system weight, which is limited by payload weight and mass balance of an airship.

2. Electrochemical generators to a certain extent meet the requirements to aeronautical complexes energy characteristics and operational capacity. However, we need to allocate a stock of fuel cells (reagents) that also results in unacceptable increase in weight of electric power system. Thus, it imposes the same limitations as chemical current sources.

3. Photovoltaic system is best suited as a basis for prospective aeronautical complexes electric power system. Photovoltaic systems, namely solar batteries, enable electric power generation directly from Solar light energy. It compares favorably with energy characteristics, high-altitude flight capability and relatively low weight. To provide on-board equipment operation in the nighttime the photovoltaic system includes battery charge-discharge controller, as well as batteries themselves. It, somehow, burdens the whole system. However, due to implementation of thin-film solar batteries the total weight of the system is not very high.

Thus, the possibility to generate electric power almost in every point of the Globe, stability of performances, high reliability, and long service life without maintenance in all climatic zones combined with absence of give-away factors include photovoltaic power system among prospective electric power sources. Photovoltaic power system combined with batteries offers necessary and sufficient electric power system for prospective aeronautical complex.

Consideration of temperature dependence of gas thermodynamic properties is necessary to solve the gas dynamics equations at high temperatures.

The effect of the variability air thermodynamic properties such as the heat capacity and the adiabatic coefficient on the Riemann problems solution was studied. The Riemann problem was solved in prediction of constant gas heat capacity in all volume, in prediction of different heat capacity from left and right of the contact discontinuity, and finally considering the real gas thermodynamics. These approaches were compared on the groundwork of three test problems with the solutions containing two shock waves, one shock wave and one depression wave, and two depression waves correspondingly. It was shown that for the Riemann problem with high temperature difference the second approach gives better results as compared with the first one, but sometimes it results some unphysical simple discontinuities appearing at points of contact discontinuities.

The effect of used numerical schemes to the solution of the Riemann problem was also examined. The Euler equation for the supersonic flow around wedge and berm was solved by Godunov scheme and its TVD modification using exact Riemann solvers for gas with constant or variable heat capacity. For all mentioned problems the computed air stagnation temperatures behind shock wave were compared. For the supersonic flow (M=10) around the berm modeled by the 1st order Godunov scheme the stagnation temperature error was less than 0.5%, and the 2nd order TVD-modification scheme resulted the 0.8% and 2% relative errors for supersonic flows with M=6 and M=10 correspondingly.

It was also shown that for weak waves the flow patterns altered insignificantly.

Finally, it can be noted that the use of Riemann problems exact solution combined with high order Godunov-type solvers can be efficient for solution of gas dynamics equations for real gases with variable heat capacities. This approach for flows with strong shock waves allows one to obtain accurately flow patterns, shock wave positions, and gas parameters behind wavefronts.

This functional defines an integral characteristic of the field [3, 4].

The linearization of temperature field within the unit cell is carried out by Rayleigh-sections method [2]. The «blurring» of the material properties is by the equality , introduced; here is the effective «action» (1) with the thermal conductivity factor .

Depending on the sequence of sections and the geometry of formed areas, i. e. sections schemes, we obtain different expressions for the effective thermal conductivity of cell.

The proposed method allows one to confirm the well-known formulae and get new ones for the relative effective thermal conductivity coefficient inhomogeneous materials (having isolated inclusions, interpenetrating components, and flat-fibrous). Also we have to establish the relationships between them, to estimate the discrepancies of values, including the most unfavorable case of very inhomogeneous environment: .

To compute numerically the effective thermal conductivity value within unit cell we formulate transient initial boundary value problem of heat transfer having no artificial restrictions adopted in approximate analytic approach.

This problem was solved for all types of unit cells of considered inhomogeneous materials on the basis of the rectangular regular grid by fractional steps method [5]. We use here an implicit difference scheme obtained by the control volumes method [6]. Due to the transient statement of the problem the pseudoviscosity method [5] was used; for the quasi-steady thermal regime onset in the unit cell we take the time at which the equality of heat flows at the end faces  is true with the given confidence level p. The distribution of the temperature field in the cell corresponding to the time being a grid function is considered as approximate numerical solution of the steady problem. Applying the general «action» procedure by equating we can compute the effective thermal conductivity coefficient as follows:

                                              (2)

Here is the base value (1), are the mean temperatures on the cell end faces that are numerically computed on the known grid , and L is a cubic cells edge length.

The effective thermal conductivity for various unit cells of inhomogeneous materials mentioned above that were computed on the groundwork of the formula (2) agree with the existing as well as new combined approximate analytic formulae.

A practical implementation was shown by computing the effective thermal conductivity of various inhomogeneous materials for aircrafts, such as the CFRP. The technique of restoring of effective thermal conductivities was developed for binary metal alloys with limited solubility of components. The efficiency of the proposed approach was shown for the binary system Al- Mg at the temperatures between 20...100...200(С. Also the effective thermal conductivities of triple-component alloy AMg5P (1557) were obtained using some different methods.

Torsional Damper and Detuner (TDD) is a mechanism for detuning and attenuation of low- frequency oscillations of OHL conductors galloping. The oscillation energy is dissipated in the damping unit (DU), which represents a structure of two disks interconnected by the joint revolution axis and an ensemble of viscous-elastic elements, which react in relative rotation of disks. One of disks is a leader and is connected to the phase conductors via a rigid frame. Another disk is a follower and is connected to the balancing lever of pendulum type.

A mathematic model of TDD functionality is developed for selection of geometric, mass and viscous- elastic parameters enabling the effective operation of the device in the frequency range 0.1-0.4 Hz. This model takes into account the hysteretic type of power dissipation appropriate to the type of DU under consideration.

The process of hysteretic deformation of viscous-elastic elements due to their displacement inside DU is of extreme complexity and cannot be subject to a rigorous mathematic description. That is why the approximate methods are traditionally employed making use of experimental data. One of approaches consists in substitution of hysteresis by viscous-elastic deformation using the tension-compression diagrams of the

elastoplastic samples and a logarithmic decrement of vibrations. Another possible approach consists in assignment a-priori of a diagram of deformation after a model of Drucker-Prager. According to this model, the hysteresis diagram is represented by a parallelogram inclined by a certain angle around its center. However, the use of such models may cause serious errors in selection of TDD structural parameters, since they do not take into account the shape and «size» of path along the deformation diagram which defines the stiffness of a damping unit and its transfer function.

In this work a cinematic model of hysteresis is suggested which relates upon the hysteresis diagram. According to it, the torque and relative angle of torsion are related one to another via a special differential equation of the first order, its coefficients being defined after experimental values for the limit hysteretic cycle. In this case one succeeds to describe, in a single equation, an infinite set of similar cycles; each of them is uniquely defined by a position of a starting point on a deformation diagram inside a limit cycle. Similarity of these curves is defined by their asymptotic approximation to the limit cycle curve. Such a model results in a natural definition of hysteretic cycle «orbital» under external harmonic action on DU.

1. Analyze features of the aviation industry and target indicators of industry development , set out in program documents.
2. Justify the need to develop an integrated educational system covering all types of professional education and all categories of employees in organizations of the industry.
3. Show the place of national research university in the implementation of construction of such education system.

Methodology of work/results

Used the methodology of constructing educational systems for various categories of students in the context of different age groups and skill levels of employers at enterprises of aviation industry.

Considered possibility of a national research university in the formation of human capital of the aviation industry. Defined the levels of final learning outcomes and their contents.

Show the ways of improving university management system, implementing an integrated system of education.

Conclusion

The article shows the need to establish an integrated system of training of human resources in aviation industry, which includes all the blocks of the educational process.

Justified expediency of ideological concentration throughout the system on a national research university, which is able to bring together, for the benefit of industry, educational institutions of various levels and education structures of industry enterprises.

Formed proposals for reorientation of university management to centers on the creation of value whic hare the department (cathedrals) that directly implement the target function in university.

Subject of study in the article are models and methods of personnel reserve formation of enterprise modernization projects. The aim is to develop methods of competency provision evaluation of the staff in order to improve the efficiency of personnel maintenance of the project.

Tasks are: planning and estimation of goods production volumes taking into account the presence of raw materials, power, ware-house and other resources, the imaginary from products realization profit and production inputs; evaluation of the required composition of enterprise equipment and the perspective plan of production, composition and technical descriptions of equipment.

Methods are: mathematical models of production plans optimization with stochastic presentation of parameters; method of indefinite multipliers of Lagrange.

The following results were obtained. It is shown that problem of production planning is in unstable terms, in particular: level of material well-being by resources, occured from realization of products, expenses on storage of unsold parts, losses from a deficit taken to the task of the mathematical programming. The exact and close algorithms are offered for the decision-making in this task. Parametrization of task, providing possibility of construction of calculable procedure decision, adaptive to changing descriptions of production and products realization, is conducted. Intercommunication between the index of products issue planning realizability and probabilistic character of production parameters is investigated.

Conclusion. In the result the model of planning of large-nomenclature production is improved in the conditions of vagueness. It can be used for solving the production planning problem for maximal realizability in the set term taking into account the estimation of the resource providing.

This article is devoted to the 100-year anniversary from the birthday of one of the most outstanding scientists and designers in the field of missile and rocket-and-space technology, Vladimir Nikolayevich Chelomey.
The area of his scientific interests is analysis which is linked with practical realization of technical systems and complexes created by him and meant to assure the countrys defence capability.
Three main areas of V.N.Chelomey activity as the designer of missile and rocket-and-space technology are identified, namely:

• weapon systems with cruise missiles,
• weapon systems with ballistic missiles,
• systems with manned and unmanned spacecraft.

Examples are given for most important technical solutions such as unfolding of missile wing during the flight which has become worldwide common practice in cruise missiles creation. Operational advantages of such solution are obvious.
A systematic approach to creation of a marine defence system is illustrated by an example of creating a sea reconnaissance and target acquisition system for the Navy. In fact, it was the first time in the world when a reconnaissance and strike system, which included systems with cruise missiles and reconnaissance spacecraft, was created.
Experience accumulated during a creation of launch vehicle designed for the sea reconnaissance and target acquisition spacecraft helped to create an intercontinental ballistic missile. Big-scale production of such ICBMs made it possible to achieve strategic parity with the USA both in quantity of ICBMs and in quantity of warheads.
Photographs show Almaz system orbital mannedstation and Proton heavy launch vehicle. Crews such as Pavel Popovich and Yury Artyukhin, Boris Volynov and Vitaly Zholobov, Victor Gorbatko and Yury Glazkov performed a lot of work on board of Almaz stations and proved correctness of design solutions adopted while creating these stations.
Time period of V.N.Chelomey formation as a scientist, a designer and a statesman is period when he was a USSR Supreme Soviet deputy.
The basic principle of V.N. Chelomey school formation is focused on the ability to put the accumulated knowledge to practical use and move it with courage to the desired goal.

The conditions that necessitated the creation and development of Navy cruise missile with special warheads aimed for ground targets destroy are analyzed.
Navy played the main role in solution to this problem by creating submarine forces armed with nuclear warhead missile weapon which can come to the shore line of USA and respond with the blow for a vitally important target. Taking into account the lack of ballistic missile in this period, the creation of submarine forces was the most effective antiweapon.
The development of cruise missiles in three designing departments of the Air Industry Ministry under the direction of G. M. Beriev, S. V. Ilyushin and V.N. Chelomey was analyzed in the presented article. Developments were virtually carried out on a competitive basis. Cruise missile P-10 was developed at the development design office named after G. M. Beriev. Cruise missile P-20 was developed at the S.V. Ilyushins development design office. Cruise missile P-5 was created at the V.N. Chelomeys development design office. There are given main tactical and operational characteristics and specifications of this missiles and development advantages of Chelomeys missiles put into service.
We have compared the cruise missile P-5 and similar to it the cruise missile «Regulus» developed in USA.
We provided the overall potential of Navy resulted from the armament of submarines with cruise missile P-5 and ballistic missiles designed for armament submarines to destroy ground targets.
The importance of cruise missile P-5 development was provided to ensure the defense potential of country in the middle of the 50s of last century.
After the development of cruise missile with a long range capability for ground targets destroy, during the last 1950 and early 1960 V. N. Chelomey had focused his great attention on the development of antiship missile (ASM) with OTH range (P-6, P-35, next was P-500) and underwater launched missiles called «Amethyst» (the first ASM with underwater launch in the world) and «Malachite», which were added to armoury in 1968 and 1972 respectively.
At the same time, in early 1960s V.N. Chelomey took the lead of the development of the system designed for space and maritime reconnaissance and target designation (SMRGD) which was added to the armoury in 1975. Submarine and surface ships with antiship missiles with a long-range capability have been equipped with ship complex of target guidance for this system. It could provide effective weapon use against ship formation with a full range in any place of World Ocean. Actually, the first effective reconnaissance-strike system was developed at that time.
Complex «Granite» which does not have any prototypes in the world and which is in operational service of Navy at the present time was added to the armoury in 1983.
The team under the direction of V. N. Chelomey restarted the development of cruise missile with a long range capability and nuclear warhead aimed to destroy ground targets in the middle 1970s (complex «Meteorite») for Navy and Air forces. However, according to conditions of political requirements, it was not added to the armoury.

One of progressive applied methods for solution to a problem of nonlinear programming in difficult technical system such as the aircraft is presented. Process of a system shape choice always takes place in conditions of incomplete information. It is caused by uncertainty both in real conditions application and prediction of preliminary technical solutions feasibility. The choice of the only optimum decision is conditional and therefore unacceptable. With respect to conditions the one decision can be difficult to realize and very expensive in comparison with other possible solutions, which earlier were seemed not optimum and therefore were not revealed by traditional methods.
To identify the set of possible extreme variants, which at further specification of conditions could become an optimum (more precisely to tell, the rational supremum by principle), it is proposed to combine the method of lexicographic multicriteria optimization with searching method with an adjustable choice for direction and size of searching steps and with fixing for partial extrema. The lexicographic component of [1] method allows to cope with difficulties of multicriteria optimization when each purpose of priority criterion leads to the alternative solutions. Uncertainty is parried by consecutive purpose of a priority row.
The problem of a choice for system design shape is solved as follows:

• some options for lexicographic sequence of criteria are formed by their priority;
• rational decisions versions of system shape for each lexicographic sequence of criteria are defined;
• the common or close shape decisions for all options of lexicographic sequence of criteria are shown.

The stated approach prepares a basis for repeated optimization.
For the repeated solution to an optimization problem the one of progressive applied methods of the solution to a problem of nonlinear programming, so-called gradient methods of a big step [2, 3], is offered. Feature of the method is that the choice of the next step direction of search is carried out with respect to the maximum gradient of function change. In the chosen direction the all partial extrema are fixed and the size of the next searching step is formed. The method allows to significantly reduce searching time in comparison with well-known methods of «formal search» and increase a number of settlement options.
Thus, the advantage of the combined method allows to optimize function with a large number of variables and many extrema with linear and nonlinear restrictions in searching area. Such opportunity is especially actual for solution to engineering tasks in choice of difficult technical shape or other (information, economic, social, etc.) systems.

This paper presents a review of the following issues for the satellite formation flying design and control problem: group configuration, reference orbit formation, intergroup subordination. Results are based on the mathematical simulation of formation flying on low earth orbits.
The term «formation flying» implies movement at least of two satellites with maximum distance between them which is much less than the length of the orbit. Satellites maintain their location with respect to the reference («central») orbit. According to the term the two types of formations can be defined: order (array) implies constant group formation; cluster («swarm») allows variable intersatellite distance.
The first well-known method for group configuration implies periodic impulse correction of each flying satellite position formation [5]. This fuel wasteful method allows array formation. Another energy efficient method provides configurations of satellites on nocoplanar orbits two times a loop (cluster formation). Other methods include either combination of methods mentioned above or exotic ways of flight formation (using electromagnetic forces [3], rigid connection of satellites and so on).
Three main methods of reference orbit formation are presented: following the leader satellite [2]; forming the reference orbit indirectly by making each satellite follow its priory defined orbit; periodic automatic estimation or the reference orbit parameters subject to current parameters of satellites and optimization criteria (minimal fuel cost, minimal time of formation, etc.).
Described control structures for formation flying satellites include: single-level structure of satellites with equal possibilities; two-level structure with a leader. The leader is authorized to control «slave» satellites (so called «total hierarchical» structure), to broadcast each «slave» personal trajectory data («hierarchical» structure) or to broadcast leader personal trajectory data («lite hierarchical» structure). Multilevel structure is introduced.

Purpose
To obtain the optimal dynamic performance of the pneumatic ground launching device, which reduce the length of the guide by upgrading expansion engine.
Design / methodology / approach
To identify the negative factors of pneumatic ground launching device workflow. To determine the path for modernization of expansion engine to get the best dynamic performance. To develop the conjugate model of catapults and optimization method aimed to reduce the length of the rail. To perform a series of experiments and carry out the results analysis with respect to the efficiency of using additional piston.
Findings
An improved scheme of expansion engine of ground pneumatic catapult, which is protected by a patent for utility model, is proposed. Comprehensive physical- mathematical model of the launching device is compiled. Method for the optimization and evaluation of the criterion of excellence dynamic characteristics is developed. The analysis of physical fields in the cylinder catapults is carried out and considerable intensity of oscillation process is shown.
Practical implications
Software package in the language Fortran 6.6 for numerical studies is created. The optimum characteristics of the proposed catapult, namely the mass of the intermediate piston and its initial position, are calculated. Synchronization move master and slave pistons are able to fix the front of the compression wave in the center of pneumatic chambers. Obtained traction law closed to the constant reduces the length of the guide by 25%.
Originality / value
Scheme of expansion engine of circuit ground launching device with an intermediate piston, which is confirmed by a patent for utility model, is firstly used. Method of dealing with the negative factors of the oscillatory process in the cylinder is first proposed. Presented range of changes in the characteristics of the proposed catapult allows to obtain optimal parameters of unmanned aircraft vehicle launch from the surface.

For achievement of hypersonic speed of pilotless flying machines it is necessary to find the solution of some technical problems connected with processes of an aerodynamic flow of flying machine surface, as well as heating and strength of its structure elements. It should be also performed the works on creation of the athodyd engine (Ramjet) with wide range regulated speed air stream. With respect of perspective multimode supersonic Ramjet working process investigation the ballistic problem of the flying machine with Ramjet (object of test) test on experimental installation «Rocket rail track» is considered. The installation represents monorail track specially designed with a small gradient angle in length of 3,5 km (nowadays the works on its lengthening up to 5 km are conducted) on which the rail carriage set powered by solid fuel rocket engines with the thrust 450 кН is placed. The increase of installation thrust-to-weight ratio level is possible by increasing of engines quantity.
In the work it is created the mathematical model of object of the test acceleration dynamics, which takes into consideration the losses: of lifting force creation, overcoming of rail — carriage boots friction forces, aerodynamic drug, vortex formation, base drug and wave resistance. The algorithm is developed on the basis of Runge-Kutta method of 4th order. The simulation code is executed using the Fortran. It were carried out different simulations with various placing of object of test and engines using one, or two rail carriages. The analysis about influence on acceleration dynamics the engines thrust reserve and variation in weight object of test or engines was conducted. The model was verified by comparison of calculation results with the obtained earlier experimental data. The executed calculation research allows to optimize the thruster and of object of test cross-section and weight for achievement of necessary speed on the 3,5 km track length. It was shown that under high density of air at speed of 3-4 M the drug losses were dominated. The tests for «Rocket rail track» installation under the 3-4 M can model the processes at the investigated flight height with a speed multiple to the air density ratio.

The article outlines the main trends of hypersonic technology development by example of carried out in the world researches. The analysis and synthesis of these problematic issues of hypersonic technology development in the field of high-speed aircraft (AC) shows:

• development of a rational integral aerodynamic layout of the airframe and propulsion system, which provides the aerodynamic quality necessary for implementing the specified performance characteristics, studies of the interaction between shock waves and boundary layer;
• development economy ramjet propulsion for cruising with hydrocarbon fuels: providing a supply and efficient fuel combustion in a supersonic flow, the introduction of new types of highly efficient hydrocarbon fuels;
• criation of effective thermal power design schemes and warm durable structure materials to work under aerodynamic heating T up to 1200 °С at the surface of the airframe at cruise speed M ≈ 6 at a height of flight about H ≈ 30 km and heating of the propulsion system (in chamber Т_с≈ 2500 °С).

• wingless aerodynamic design with a monocoque body and the air intake with a sub-screen;
• propulsion: solid rocket fuel (accelerators) and marching ramjet hydrocarbon fuels;
• thermally loaded airframe components (T > 1000 °С) and propulsion system of high-temperature composite materials «carbon-carbon», «carbon- silicon carbide» with antioxidant nano coatings for less thermally loaded structures (T = 600 ÷ 1000 °С) — high-temperature titanium, nickel alloys, niobium-based, heat-resistant coatings and special active cooling system.

In the world is given the highest priority to backlog development in scientific, technological, methodological aspects for hypersonic aircraft. In the next decade on this basis will be created high-speed devices for different applications. Among the leading foreign countries the most progress in the implementation have made of U.S. government programs Hyper-X, Hy Tech, X-51A, Hy Fly etc.

It is proposed to use crash-boxes instead of traditional honeycomb elements in the spacecraft landing shock-absorbing device. The developed design- parameter selection procedure provides a safe spacecraft landing process.
The spacecraft is considered as a point mass equipped with a single support contacting the ground. The formulated problem of non-stationary dynamics is solved using the finite element method, the explicit integration procedure, and the combined MSC.Dytran- MSC.Patran software.
It is assumed, that this single support may include one or more branches of energy-absorbers. Each branch is a group of crash-boxes connected in series by support rings. Axes of the same-branch crash-boxes are aligned. There are three structural schemes of the considered landing device: 1) the single-branch device, 2) the device made of two parallel branches, 3) the device made of two non-parallel branches.
Kinematic and dynamic transients of the specified structural schemes are integrated till the spacecraft point- mass unit overload is obtained. The simulation time- dependent curves characterize the spacecraft landing velocity and acceleration. It is shown, that the point- pass overload level provided by the crash-boxes is close to the overload level provided by the traditional honeycomb elements. Advantage of the proposed crush-box support is its effective damping the peak shock loads.
It is shown, that non-availability of plastic hinges in the absorbing device is the necessary condition of stability of the spacecraft landing process. Longitudinal shock conditions are the most favorable ones for effective shock-absorbing by a separate crush-box. The proposed new non-parallel structural scheme provides the specified longitudinal-shock landing mode.
The obtained simulation results show, that the yielded crash-boxes can provide up to 40% mass saving of the landing-device. These practices may be used by spacecraft engineers for developing future landing devices of re-entering vehicles and interplanetary space-stations.

Ballistic analysis of the four-revolution scheme of the rendezvous between transport vehicles (TV) Progress and Soyuz and the International Space Station (ISS) was carried out.
In the case of four-revolution scheme the ISS/Earth/ TV phase angle stays in the range of 20-40 degrees by the time of rendezvous orbit generation. For this purpose, before TV is launched ISS working orbit is built by means of a few maneuvers. The latest ISS orbit correction is performed about 1-2 weeks before TV launch in order to minimize the atmospheric model errors, geomagnetic and solar activity fluctuations and the space debris collision probability.
The first two impulses which provide for the required rendezvous conditions are implemented after TV launch between the 1st and 2nd orbits beyond Russian ground facilities visibility zone. These impulses are calculated using the nominal entry conditions 1 day before launch and are sent to TV prior to launch.
Due to launch errors the 3 rd and 4 th impulses are implemented on the 2nd and 3rd orbits, in order to compensate for revolution period T, LAN Ω and inclination errors. According to the statistics of 45 launches within the period of 04/07/2007 through 10/ 31/2012 the actual errors were as follows: ó_T= (5.161+3.736) s, óΩ = (1.1+0.7)·10-3 radian, ó_i= (3.6+1.8)·10-4 radian.
The satellite navigation system onboard TV allows to form the required orbit without the 3rd and 4th impulses. Flight tests of this system are planned to be performed in 2015 onboard Progress M-26.
The 3rd and 4th correction burns are calculated onboard basing on the state vector obtained by ground stations at the beginning of the 1st orbit and on the magnitudes of the 1st and 2nd burns. TV and ISS state vectors and information about solar and geomagnetic activity indexes are uplinked in the visible zone of the 2nd orbit.
Final decision about the autonomous approach or transition to the two days rendezvous is made basing on the telemetry data of the 3rd orbit.
Since there is no onboard navigation system, the orbit determination accuracy is calculated basing on the KAMA measuring system information. A priori and a posteriori calculations show that the errors in aiming point, selected on the 4th orbit, do not exceed 80 km by 3 ó along the orbit.
Successful docking of 5 Progress and 4 Soyuz vehicles allow to make a conclusion that the 4th orbit rendezvous scheme can be adopted as a nominal one.

Statement of research problem
Multilevel optimization scheme aircraft development process includes two tasks. On the top i-first level solves the problem of optimization of process of formation of system LA, specifies requirements for the reliability of subsystems and distribution of costs and development time, the establishment and operation. At the bottom i — th level optimization problem is formulated and the process of developing the aircraft reliability of its elements.
These detailed analysis of the development process for i - th (lower) level of management development take into account peculiarities of the project (technical and technological constraints). When implementing a two- tier scheme of project management are used to clarify (adaptation) project dependencies on the top i-first level. Thus, when conducting a coordinated process optimization project of LA (at the i-1st level) and process development aircraft (for i - th level) than the above direct links in the development of multi-level governance model implemented adaptation of the top- level test data models for the lower level.
The paper presents a coherent statement of the problem of two-level optimization of the development process and ensure the reliability of the aircraft and the algorithm for its solution.
Research technique
To solve the problem stated above is used statistical method agreed a two-level optimization. Implementing a multi-level optimization process of formation and reliability of aircraft to reduce the number of variable variables dealing with the related extremal problems at every level of development management, simplifies the search for optimal solutions.
Research results
The problem of optimizing the process of formation and system reliability is formulated as a multi-criteria, multi-parameter. The main criterion - the total project costs, which depend on the reliability of subsystems during the design, creation, the duration and scope of testing stages when developing. Are calculated according to the appropriate definition of the function changes the reliability and distribution costs allocated stages of the project.
The features of the process of formation and implementation of system reliability LA. The optimal control of the development of requirements for the reliability of subsystems at the end of the design phase and in the late phase of not the same. Load increase, for example, the duration of operation leads to an increase in the reliability requirements, and the necessary reliability gains for selected elements are different. Obviously, due to different points marked «cost of reliability» for subsystems RK. Results of a comprehensive study of the process of development in the presence of restrictions on the duration of the stage creating a show that a decrease in the duration of creating the project costs are rising, due to a flaw and the relatively low level of reliability of the system, the need to increase the number of aircraft for the job. These costs are also rising with increasing duration more. Calculations show that while reducing development time cost effective additional cost of ground testing stage, to increase reliability and to study the conditions of «life» of the next steps.

The existing air heat exchangers made of aluminum or stainless steel, do not allow to implement highly available power of aircraft without losing mass perfection. Thats why the leading aircraft manufacturers conduct intensive researches aimed at creating high heat exchangers made of nonmetallic materials. Composite materials can be regarded as the most attractive for the manufacture of structural elements such heat exchangers. The replacement of metals by high temperature oxidation resistant and erosion resistant composite materials, which have significantly higher levels of specific strength (relation of strength to specific gravity ) at temperatures of 700-900oC and above, will dramatically improve the mass-energy characteristics and increase the competitiveness of the new generation of products. At the same time, composite materials with the reinforcing framework made of thin nonwoven carbon materials or of silicon carbide ones produced by using of a special hydrodynamic machining (Spunlace-method) are more promising than materials derived from the tissue.
The paper presents the analysis of existing technologies of producing high-temperature heat exchangers. The technology of heat exchangers elements production from oxidation resistant composite when each element has a strength reinforcing frame, and frame filling matrix has oxidation resistant properties, seems to be the most reasonable one.
The analysis of existing basic materials and the limits of their applicability for the manufacture of high heat exchangers is fulfilled.
The paper presents the technology for producing a composite material with a silicon carbide matrix on the basis of nonwoven carbon materials for corrugation, separating plates and clamps of high temperature heat exchanger.
The results of corrugations elements saturation with oxidation resistant matrix based on silicon carbide are presented.
The results of X-ray and microstructural analysis of thin-walled composite samples by electron scanning microscopy are shown.
The parameters of corrugations and separator plates saturation process with matrix based on silicon carbide are defined. These parameters provide the volume saturation of the heat exchanger structural elements within 43-46 % of weight.
The results of practicing combining the elements of high-temperature heat exchanger into solid cartridge using high-temperature adhesive and the following assembly compaction with silicon carbide from the gas phase of ecological organometallic reagent — methylsilane CH3SiH3 are presented.

The purpose of this work is to analyze existing systems for thermo magnetic energy devices for wide spectrum of application and to describe the optimal magnetic system for thermomagnetic engines (TME).

Design/methodology/approach

The revue about magnetic systems for thermomagnetic energy devices based on permanent magnets (magnetic cooling machine, magnetic thermal generators and thermomagnetic engine) was done. The publications for the last 5 years were considered. It was proposed to classify the magnetic systems used in thermomagnetic energy conversion technologies in three groups.

Results

For magneto caloric refrigerators and thermo magnetic engine its magnetic systems have a fundamental difference in the distribution of magnetic induction along the active working ferromagnetic body.

For refrigerators magnetic systems requirements are:

• to achieve the maximum possible values of themagnetic induction in the gap of the system;
• obtaining a uniform magnetic field (grad(B/x) =0).
• Systems for thermomagnetic engine meets the following requirements:
• ensuring high peak values of magnetic induction in the working volume of the magnetic system ( ≥ 0,6 T);
• creation of a long and continuous growth of the magnetic induction to ensure acceleration of working ferromagnetic elements;
• ensuring slump drop of the magnetic induction values in the distribution of the magnetic field at the output of MS to reduce the duration of forces, braking rotor.

Research limitations/implications

This research provides development of non-traditional methods of mechanical or electrical energy production based on thermomagnetic effects.

Originality/value

There is a lack of magnetic system designing data for thermomagnetic energy devices described in literature and, on the contrary, there are a number of works in which presents an analysis of the existing systems for magnetic refrigeration. The magnetic refrigeration systems are introduced different criteria for evaluating the effectiveness and feasibility of the application of a system [15] and examines possible field configuration and geometry of magnetic systems [4, 510].

Purpose
When in service on various orbits in the course of and after magnetic storms differential charging of spacecraft structure members takes place. As a result ASD occur. Effecting external shields ESDs generate conductive EMI in inner conductors of spacecraft on- board cables leading to malfunctioning and failures in operation of spacecraft equipment. Due to its significant length the spacecraft on-board cable network is main receptor on exposure to EMI generated by ESD.
Therefore, to provide required noise immunity of spacecraft electronic units and equipment of space vehicles, it is necessary to know the levels of currents and voltages in on-board cable network on exposure to EMI generated by ESD. The accuracy of calculations depends on selection of an adequate model of ESD impact on spacecraft on-board cable network.
The goal of the paper is design and implementation of conductive electromagnetic interferences in spacecraft on-board cable network calculation algorithm using MATLAB.
Design/methodology/approach
A theoretical part is based on classical electrodynamics and electrical engineering. Also elements of complex variable theory and operational calculus were used.
Findings
The paper presents mathematical models of electrostatic discharge impact on external shield, shielded and non-shielded spacecraft on-board cable twisted pairs of wires. Algorithm for calculating conductive electromagnetic interferences in spacecraft on-board cables, generated by electrostatic discharges, is described. An example of a pulse voltage waveform over the inner conductor of non-shielded twisted pair obtained with MATLAB is presented. research.
Research limitations/implications
The present paper provides a part of spacecraft on-board cables electromagnetic interferences calculation research.

When an aircraft approaches Moscow terminal control area, an air traffic controller instructs the crew to push down and vector by beacons. Orienting by radio beacons and following air traffic controllers instructions the plain approaches a final approach track or a landing trace matching strip magnetic direction. Selection of the strip itself depends on a number of factors, such as saving and flight safety account under limited fuel reserve, and performed according to a special algorithm.
The problem of airplane direction to one of several runways at different airfields is considered. Initial landing heading angles are set for every runway. These angles determine final courses. The total number of runways is also specified. This number defines wind heading angle which value may be changed. Depending on its direction one of two final courses is sel ected for each runway. The problem of aircraft direction during safe landing approach is considered for each runway. Horizontal flight is analyzed herewith at a given constant height. Each aircraft is characterized current time by state vector defined by the following coordinates: the shortest distance from an aircraft to a course line, minimal distance to the nearest aircraft at the flight level already on the given course line, heading angle reckoned with respect to a course line, fuel reserve spent for additional maneuvering. Parameters, accepted as constant are as follows: known flight speed, maximum allowable lateral acceleration during turns, minimal safety traffic distance at the flight level. With a shortage of space on the trace for aircraft safe traffic, a part of them is directed to a queue of this track, called «trombone», so that later land at the same airport at the earliest opportunity. Each of these aircraft current state coordinates is changeable in accordance with the known differential equations of motion describing the flight dynamics. For simplicity, each coordinate thereby corresponds to a single differential equation. One of the most important assumptions is selection of the air traffic control optimality integral criterion, which should estimate the flight efficiency in convolution. In this work, we adopt such criterion as minimum integral functional, which accounts for undesirable penalty deviation fr om the track, runway course, the distance between adjacent aircrafts on the track itself and the amount of spent fuel reserve. We decided herein to take the path of unchangeable attainment of the preset guaranteed safe movement distances between aircrafts, and this is of paramount importance.
It is required to determine the final runway courses with provision for wind direction, and allocate aircrafts among traces. Doing this requires generation of the table of priorities for all aircrafts. On the ground of this table that lists aircrafts for each runway, it is necessary to determine the priority of the approach and landing of aircrafts for each trace.
The author developed an algorithm for final course auto detection with provision for wind direction. Also solved the problem of dynamic priorities computation of a commercial airplane entering preset flight level or formation. One should take into account the possibility of this priority increase herewith. The author offers an automated solution to the problem of forming the sequence of arrival priority for randomly positioned in space aircrafts and their entering the standard arrival route.

The article is devoted to an actual problem from a field of strength of aviation and rocket-cosmic thin- walled constructions such as solution to boundary problems for a definition of their stress-deformed state on the basis of high defining differential equations in the partial derivatives. An object for investigations is the circle cylindrical shells made from material, which has different elastic properties in the three mutually perpendicular directions, i.e. the physically orthotropic cylindrical shells.
Under the action of arbitrarily changing on shell surface circle loads in the article for the first time there were obtained differential equations of the eighth order in the partial derivatives of general theory for the physically orthotropic shells most often meeting among anisotropic thin-walled constructions. From sufficiently great variety of initial equalities of the shells theory the authors sel ected the Vlassovs version as the most coherent fr om the point of view of the energy-static principles compliance. In particular, the Betti works return theorem is exactly executed, that causes the symmetry of the three equations system in displacements for the theory of shells. Thus, in the article the Vlassovs theory of isotropic cylindrical shells is generalized to the case of the physically orthotropic shells, i.e. the new differential equations are obtained.
Because the integration of the differential equations of eighth order in the partial derivatives of the general theory for the physically orthotropic shells is a great problem, a simplification is made in the article on the basis of confrontation principle of the stress-deformed states variability in axis and circle directions (Novogilovs criterion). As a result there were obtained the new differential equations in the partial derivatives of «gently sloping shells type», which for the isotropic shells transient to the Vlassov-Donnells equations; «semimomentless theory type » for the isotropic shells transient to the widely well-known equations; and the equations for the tangential stress-deformed state with very high variability, i.e. the analogue to the plane problem of the elasticity theory.
On the basis of obtained differential equations the boundary problems for the finite long closed cylindrical shell with hinge fixed edges under action of circle local loads are considered. The differential equations of the general theory are applied for solution to this problem by Fourier trigonometric series method.
Some results of calculation are represented in the graphic form showing the influence of loading type and orthotropic parameter on the stress-deformed state of shells.

The nonlinear plane contact problem of indenting a hard punch in an elastic body with functionally graded coating is considered. The main vector and the main moment of the external forces applied to a hard punch are assumed set. Unilateral contact conditions are given on potential contact area on which the elastic body and the hard punch may come in contact. Thickness of the coating is commensurable with the sizes of the spot of contact. Such scheme is applied at research of discrete contact of rough bodies.
The variational method is used to solve the problem. The variational formulation of the problem in the form of a variational inequality and an extreme problem areobtained. The variational inequality is the formulation of the variational principles of virtual work and the extreme problem is the formulation of the variational principles of minimum potential energy.
The discretization of the problem is made by an unstructured triangular finite element grid of the axisymmetric finite elements with triangular cross section. The stiffness matrix of the elastic body is a sparse block matrix. The block sparse row format is used for saving it. This scheme of storage of the stiffness matrix imposes minimum requirements to memory, which volume is proportional to the first degree of number of unknown, and allows to organize effectively calculation matrix-vector multiplication. Modifications of the conjugate gradients method is used for numerical solving of the obtained finite-dimensional nonlinear programming problems.
Developed computational algorithms are implemented as a separate module of an application package for solution of non-homogeneous elastic finite bodies contact problems.
The carried-out calculations showed that existence of functionally graded coating, which elasticity module exceeds the elasticity module of the main material, leads the maximum values of contact pressure to increase the maximum values of contact pressure and the punch penetration to reduce.

The subject of research in this work is models and methods of forming of skilled reserve of enterprise modernization projects. The aim of work is development of estimation methods of employees competent reserve for the increase of efficiency of the skilled providing of project.
Decided tasks are followings: to ground importance of estimation of employees professional competence at forming of skilled reserve, to develop the method of estimation of competent reserve of engineer-technical workers, to make an example of determination of employees optimum composition.
Applied methods are: theory of sets and boolean functions; methods of discrete optimization, in particular, integer programming.
Such results were got. Basic functions and expected results of creation of the system of skilled competent reserve are certain. It is indicated that errors at forming and management skilled reserve of project reduce efficiency of his implementation and can become reason of derangement or waiver of works implementation, and also are reason of considerable additional expenses. Factors and features of labour resources which must be taken into account at creation of the system of employees backuping are analysed. It is offered formalization of the indicated task with the use of sets theory.
The method of estimation of competent reserve of engineer-technical workers is created. The mathematical model of forming of reserve is developed as a task about the least coverage of sets. The mathematical theory of discrete optimization is used for the decision of task about vertex coverage on graphs.
It is indicated that the existent mathematical methods of decision of tasks about finding of coverage do not take into account the specific of management skilled reserve. In this connection, the method of decision of task of employees reserve forming is improved. Substantive provisions and stages of method are described. It is indicated that a method allows to get the set of possible decisions and choose the best decision on the criterion of minimization of cost.
Conclusion. The models and methods offered in- process can be used for development of algorithms and information technologies of management the skilled providing of projects, including at forming of skilled reserve of project of enterprise development.

The subject of this article is the statement of the problem of creation of a new organizational-economic concept of the defense article research and development contract management system that take into consideration the features of research and development and that provide the recovery of the research-and-production potential of defense article developers and manufacturers and that corresponds with the idea of the partnership between private company and state. The article shows that the new organizational-economic concept of the defense article research and development contract management system should be based on the new organizational-economic approaches of the state customer to the setting of the pricing condition, funding conditions and production-property conditions of defense article development beyond the legal space of work and labour.
The management of the new defense articles is based on the existing defense article research and development contract management system and development process requirements.
The state customer determines and fixes the development price limit. Engineering risks connected with the new knowledge created during the development are typical for research and advanced development.
Realization of the engineering risks fairly leads to the of the term and rise of the research and development costs. The existing methods of development costs prediction dont let the price raise be considered.
During the research and advanced development in the sphere of science intensive defense article the following features take place (towards the main contractor of the research and advanced development):

• strict fixation of the term and funding size of the research and advanced development,
• joint funding of the research and advanced development,
• mix (budget and non-budget) funding of the research and advanced development.

The possible term and funding risks of the general contractor is not taking into consideration. Strict fixation of the term and costs of the research and advanced development steps makes the contract development management system inflexible.
Joint and mix funding of the research and advanced development are realized in practice without the necessary normative-methodical documents.
Contract development management system of the research and advanced development doesnt consider production-property conditions typical for the development.
Consideration of the development features in current legislation and reflection of these features in interdepartmental (departmental) normative-methodical documents on development is possible on the basis of a new science-based organization-economic concept of development management contract system that should consider development features and provide recovery of scientific-production potential of the developers that was lost during the domestic economy reforms.

Proposed and implemented in practice the scientific- methodical approach to work to develop a Comprehensive plan of research in the part of the state aviation in the framework of the National development plan of science and technology in the aviation industry until 2025 and the subsequent perspective (further the National plan) allowed joint decision representatives of enterprises and organizations of the aviation industry, theResearch Organizations of theRussian Defense Ministry:

• to clarify and adjust bids to perform research;
  
• linked to the development of research claimed the final products of the government programs, primarily to assess their main indicators to conclude that the consistency of these studies and decide whether further consideration;
  
• expedite and facilitate the decision-making experts on the priority of particular studies and their inclusion in the Comprehensive Plan of research, given the established limits funding for the National Plan.

Funding research and development (R&D) in aerospace industry is associated with such uncertainties as the value of work and the achievement of specified technical, technological, financial and economic results. Securitization could be used as an insurance method. Securitization is assets funding by means of issuance of derivative securities based on some pool of assets.
There are different types of securitization. Traditional securitization implies that illiquid assets or liabilities are converted into liquid derivatives and their actual sale takes place. Its purpose is to provide cheap funding sources. Synthetic securitization is an issue of instruments which execution is determined by event specified in the underlying asset. In this situation collateral transaction is realized instead of selling. The main function is to hedge risks.
It is suggested to use securitization methodology in determining the value of R & D in the aerospace industry. To ensure traded derivatives not only underlying assets but also liabilities can be used. As the assets and liabilities may be characteristic Development stages of R & D, some work and unfinished study, volume of the funding needed to carry out R & D and other assets could be used assets and liabilities to security traded derivatives.
During the securitization of research and development in the aerospace industry it is suggested to use different types of contracts being derivative financial ’instruments: options, futures, exchange forwards, OTC forwards and swaps. Considered derivative characteristics assumes the possibility not only to securitize financial obligations, but also obligations for conducting the research and performance of work during the R & D. Specified technical indicators, complexity and cost of research could be used as underlying asset. Securitization allows distributing the possible risks between the parties as the liabilities could be transferred as a whole and as a part. When the part of work is transferred to contractor R & D value could be defined as the specified amount in the contract instead of calculation the costs of the certain part of work. Thus the use of the securitization methodology in the organization of contractual relations during research and development and calculation of its value is suggested.

The paper considers the ways of increasing the efficiency of lifting area of the soft wings, which are used in the gliding descent and powered flight systems, via engineering solutions. These approaches are based on making the wing rotate about the horizontal axis of the craft and thus create a rotation lifting surface. This way makes it possible to reduce required wing area by more than twice. At that the distinctive features of the «rotoplan» configuration, which is created on the basis of a soft wing, would allow the designer to avoid the known drawbacks of the existing cyclogiro configurations.
The known cyclogyro configurations use multibladed rotors, which are installed on the bearing axes. Large aircraft, which are built according to this configuration, prove to not be efficient. This is caused by excessive weight of rotation axle and blades, which carry a high bending load. Another reason is the additional drag of the structural elements, which support the blades. The ratio of the disc area to the blades area is also worse for a cyclogyro than for a helicopter. However, the use of the soft wing rotation can help to avoid the described cyclogyro disadvantages. In particular:

1. the wing area can bequite large without aserious increase inthe weight;
   
2. the wing itself isinastretched state during the rotation. Atthat itissupported bylines and thus itcan withstand the rotation about the horizontal axis without any additional structural elements;
   
3. norotation axles are required tosupport the soft rotating wing;
   
4. it is possible tochange the shape ofthe wing and its setting angle byusing the available lines without the need for any additional structural elements.

During the flight in the process of separation of aircraft moving mass as a rule the high dynamic loads appear on the aircraft, its equipment and construction elements of small systems on board. As an avionics on aircraft board can be instrumentation equipment, control systems and small-sized engines. As a result of the flight and ground experiments the question of received data processing with unconventional ways arises. In this paper the construction of empirical laws of vibration loads distribution is proposed by using high-torque statistics.
It is shown that the considered spectral array of samples shares into vertical columns «tied» to a fixed frequency range or to the narrow intervals around these frequencies, for example, to the 1/3-octave frequency bands. In each vertical columns of described spectral model the spectral vibration statistic is formed from sample spectrum of certain amount that qualifies as a random variable.
In this paper it is proposed to estimate the experimental results and the law of distribution in terms of asymmetry and excess. A distinctive feature of the presented solution for moments order more than two is that this approach for dealing with the empirical distribution provides much better rate of convergence in comparison with the similar solutions using low-torque vibration statistic when the moments order in less or equal to two.
The paper presents a table representation of both solutions. The result is a generalization of the probabilistic model of random vibration loads spectral characteristics for certification assessment of the vibration environment on the aircraft.
In conclusion we can draw the following statements: from the available data with the use of high-torque statistics it is possible to obtain distribution law characteristics enabling widespread in current processes evaluating.

Magnetic fluid (MF) technical means have very important property — they are controlled, so they spread in pneumatic, hydraulic, electric, electronic and other types of technics as a working body, in which the magnetic liquid changing the rheological properties (magnetization, viscosity and density is applied; electric, optical, acoustic, etc.) under the influence of an external magnetic field is used. Now three main types of MF are known: ferrofluids, magnetorheological suspensions and magnetic compositions.
Now the main directions of application of MF is the following: 1) magnetic fluid systems of vibroand shock defense; 2) magnetic liquid consolidations; 3) magnetic liquid additives for fuels of internal combustion engines; 4) magnetic liquid fluid transmission and transmissions of vehicles; 5) magnetic liquid greasing; 6) magnetic liquid water purification from oil products; 7) magnetic liquid separation of not magnetic particles; 8) magnetic liquid support and bearings; 9) magnetic liquid measuring devices; 10) magnetic liquid heat carriers; 11) magnetic liquid chemical technologies; 12) magnetic liquid printers; 13) magnetic liquid distributive equipment for working bodies (working liquids, gases); 14) magnetic liquid medical devices.
A number of tasks were solved during the creation of movement speed automatic control system of the mobile unit with application of automatic magnetic liquid transmission for the mobile units (MU) of missile and space-rocket systems. Application estimation of MU movement control system with automatic hydrodynamic transmission based on the operated magnetic liquid fluid transmission (FT) is carried out.
Thus, MFwith ability to change their sizes and properties by the meaning of influence of an external magnetic field are fully belonged to micro and nanomaterials, and their use in systems of vibroprotectionof heavy-loaded and light cars, buildings and constructions in systems of branch and transformation of heat of space-rocket complexes units will allow tosignificantly improve their operational and technical characteristics.
Further it is reasonable to use magnetic fluid systems in creation of new and modernization of existing spring systems, systems of vibroand shock defense, transmissions, systems of vehiclesgreasing, systems of cars safety at the industry enterprises — BAZ, MZKT, KZKT, Ural Automobile Plant, KAMAZ, — systems of ventilation and conditioning of buildings and constructions, systems of branch and transformation of heat of various units of space-rocket complexes.

Penetrator is an impactive ingrained probe, which is placed into the soil and used for study on internal structure of celestial bodies. The peculiarity of the concerned penetrator is a solid-fuel propulsion system with a straight-through channel, which allows a rational mode of penetration by reducing the resistance forces arising in the ground.
The special features of the functioning propulsion system with a straight-through channel are soil inside intake, its mixing with fuel combustion products, vortex and outflow of the resulting mixture from the through channel.
Intake soil and mixing it with the combustion products leads to the necessity of study on two-phase flow in the chamber of variable section. Analysis of dependencies in the article shows that the friction loss of the gas particles will be increased because of intaken soil particle size growth and their concentration. Emanating combustion products volume from the nozzle of the penetrator with a straight-through channel rises in 1,5...3 times because of intaken soil on board and leads to an increase in draft at 20...30 %.
Besides these relations can be used for calculation of the rotating mixture basic parameters in the cross section of the through channel including in the critical section of the nozzle.
Thus, during the selection of optimal variant for constructing penetrator with a straight-through channel it is necessary to add the two-phase mixture degree spin and a hole diameter, through which soil is taken inside the machine, i.e. diameter of front end section of the through channel in the range of main design parameters that uniquely identify the penetrator.

At the present moment the basis for automatic system designed for preventing dangerous situations in outer space (Russian acronym «ASPOS OKP») is created in MCC TSNIIMASH under the aegis of Roskosmos. The main task of the system is collision warning between spacecraft and space objects (SO).
Collision prevent between spacecraft and space object is more economical profitable then to remove its consequences [1].
The ballistic analysis of the Russian radar stations and optic-electronic systems contribution into the process of the different SO orbit determination is in this article.
Calculations show that accuracy of SO trajectory parameters determination by the Russian space monitoring system is not enough.
The main way to solve this problem is to develop the new radar stations based on the large antennas such as OKB MEI (Cobalt radar station) [7, 8] and Ysuriysk, which can track space objects using target designation from the present radar stations. In this case the low earth SO orbit determination accuracy becomes about 50
100 m and prediction errors in the 8 hours time interval dont exceed 300500 m.
The network of special optic devices, which are the part of the segment ASPOS OKP IPM RAN, was created to provide the ASPOS OKP central core with the information about space objects high apogee. The network covers the area of geosynchronous orbits between 20° west longitude and 170° east longitude, and the area of high apogee orbits. Moreover, Roscosmos designs its own optoelectronic devices network, which can provide the accuracy 0.51’’ for GEO and 12’’ for high apogee orbits using the ASPOS OKP segment target designations. Collaborative work between optoelectronic devices and big antennas radar stations allows ASPOS OKP to estimate the collision risks for high apogee spacecraft, because the determination accuracy and orbit prediction dont exceed 400 m.

This article is devoted to the decryption of the Correction Propulsion System (CPS) named «ЭОЛ- 2010» developed on the base of the low power stationary plasma thrusters (SPT). This device is dedicated to power the low mass spacecraft keeping its working orbit parameters during active lifetime and de-orbiting after expiration. The CPS relevance is dedicated by increasing of the micro-satellites application in the various fields of science.
This CPS is designed as a monoblock structure, which adapted to install on the unified spacecraft chassis. CPS includes the propellant storage unit, propellant flow control unit, and the thruster unit. The design with both single and two thrusters modifications are described. The working mode of the thrusters was choosing with adjustable power level in the range from 150 to 210 Watts, taking into account solar cell illumination change or its degradation. The propellant storage unit tank contains 4.1 kg of xenon at a pressure of 10 MPa. The total system thrust impulse is 44 kN under full propellant utilization. The CPS dry mass is 8 kg.
The calculations presented in the article demonstrate the application possibility of micro-thrusters correcting propulsion system usage for orbital transfer and adjustments in low orbits (up to 600 km) the spacecraft mass up to 100 kg, taking into account its aerodynamic drag. It is shown that spacecraft lifetime increases 5 to 10 times with such CPS using.
The results of firing test of the thrusters aboard the CPS module with the flow rate characteristics controlled by the weight method are described. The possibility of using the CPS module as test stand equipment for testing of SPT modifications and elements of a Propellant Storage and Power Supply System is shown herein.

This work is devoted to the noise problem of light propeller aircrafts. In the article the experimental obtained basic results of acoustical characteristics by light propeller aircraft (LPA) of Vilga-35A in flying conditions are presented. Spectral and power characteristics of LPA power plant (PP), acoustical radiation and also directional characteristics are recorded.
The problem due to decrease noise of PP is rather actual for a general aviation LPA, whereas ICAO norms act on maximum permissible noise levels of such planes [1]. The given work follows of author previous researches [2] about acoustic characteristics of LPA power plants.
Measurements of LPA Vilga-35A acoustic characteristics have been carried out. The plane PP involves a piston engine (ICE) AI-14RA with 2-bladed propeller of variable pitches US-122000. Measurements of acoustic characteristics were conducted at various level flight conditions, which correspond to five power modes.
It was obtained that the acoustic field of a power plant of a LAN is determined by acoustic radiation of a propeller. The regression equations establishing the relation of aircraft noise levels on the ground from power of a PP and Mach number of relative flow velocity in tip section by propeller blade are received. It is find out that acoustical efficiency of Vilga-35A light propeller aircraft power plant on maximum power operation is 0,123 %.
Broadband radiation in the field of 1600-3000 Hz has an aerodynamic origin, and its dominating source is the turbulent wake behind propeller blades. The peak of broadband radiation in the field of 5000-8000 Hz most probably caused by radiation of a mechanical origin from planetary gearbox of the engine.
The complex approach consisting in decrease of power plant noise, i.e. the noise radiated by a propeller and the piston engine is necessary for essential decrease in noise level of LPA with ICE on the ground. Decrease of a propeller noise is enough challenge. The solution of this problem is based on optimization of geometrical, aerodynamic and acoustic characteristics of propeller. Intensity of propeller acoustical radiation under condition of conservation of constant thrust can be reduced by means of optimization ratio between number of blades, propeller diameter and circular velocity by criterion of the minimum acoustical radiation power. For decrease in the noise radiated by the piston engine, it is necessary to apply ICE exhaust silencer.

The experience of production of sandwich panels from titanium alloys via a method of superplasticforming and diffusionbonding (SPF/DB) shows that in the process of forming under certain conditions there are defects. The defects arise as a result of bending sheet of panel in excess of certain critical value. The bend of a sheet of the panel in the process of forming occurs in a free zone of the bottom sheet between diffusive welding of the next edges of the panel. The size of a bend of a forming in a free zone depends on a material of the panel, a mode of forming and a design of a filler panel.
In the work the values of critical height of a bend were determined by a method of modeling of process in the program «MSC Marc» depending on a design data of filler. Thanks to the received values the range of a choice of optimum constructive and geometrical parameters of sandwich panels was determined. The range is presented in the form of a plot for values thickness of sheet, width of the filler and panel height. Each plot shows that above the curve there is the recommended range of choices thickness sheet and filler for the forming of a sandwich panel of titanium alloy desired geometry without the defects.
Process is presented for a modeling and forming of the sandwich panel with the parameters chosen from the recommended range which confirmed a possibility of forming without manifestation of defects. To check of results of modeling the forming was done for the sandwich panel with the received parameters on the press of superplastic forming and diffusion bonding «FSP60T», in a laboratory of IRGTU. As a result of the forming the panel is received without any defects.
Thus, the results of modeling and natural production showed that when the recommendations, presented in this paper, are followed, there will be no defects in the process of forming of panels.
If its not possible to select the optimal structural and geometrical parameters of sandwich panels presented in the work, then its necessary to resort to special methods of forming, one of which is the use of supporting of sheets during the forming process to prevent the formation of bend and, as a consequence of defects.

The article investigates the possibility of modifying the surface to form a nanostructured layers and coatings in metal materials and contains some specific examples of successful implementation of the objectives of nanostructuring of the surface layer in a metal surface by engineering techniques using various sources of thermal, chemical, heat treatment, physical or combined effects. In this investigation the possibility of obtaining the dispersion-hardened layers or coatings for forming a three-dimensional type of nanocrystal is realized: namely, in micro-sized (thickness of the order of tens or hundreds of microns) layer structure of a matrix solution with dispersed nanosized precipitates of secondary phases is formed. Carbides, nitrides and borides are such nanosized phases in our scientific paper.
There were identified the following promising surface modification processes, producing nanostructure in the surface layer of metallic materials:

• processes of diffusion metallization from coating using low-temperature plasma in combination with nitriding ;
• the process of laser — radiation boriding from coatings ;
• processes of receiving pyrolytic chromium coatings by vapor deposition of the metaloorganic chromium compounds in combination with a laser annealing.

Such combined surface engineering techniques involve a combination of thermal effects (chemical heat treatment, annealing), chemical exposure (doping, coating deposition) and physical effects (plasma, laser treatment).
The considered combine processes of surface nano- modification give the possibility to achieve carbon and high-alloy layers on the low-alloy steels with elements of the nanostructure, which provides high physical and mechanical properties of materials that improve the reliability and durability of machine parts and tools.

The analysis of effect of a random impulse noise on characteristics of space communication timing recovery systems is carried out. With usage of simulation models, characteristics of statistical equivalents of timing recovery discriminators in the conditions of joint effect AWGN and a random impulse noise are determined. The received results can be used for design and the analysis of the noise immunity property of the space communication systems functioning at effect of the given aspect of pulse noise.
Since Gaussian noise and pulse noise affect both on information transmission channel and on timing recovery channels, timing errors lead to additional lowering of noise immunity and should be in advance considered at design step. In the present paper, mathematical and simulation models of timing recovery systems, their basic characteristics in the conditions of joint effect of above-mentioned noises will be observed. The greatest distribution timing recovery systems are observed which are Early-Late Gate Timing Recovery (ELG-algorithm), Gardner Timing Recovery (G- algorithm), Mueller-Muller Timing Recovery (М&M).
The presence of impulse noise and AWGN makes different impact on operation quality of different timing recovery systems. It was found that the most susceptible to impulse noise should be recognized the algorithm М&M. G-algorithm and ELG-algorithm are rather more resistant to impulse noise effect.
The obtained results can be used at designing for the selection of better parameters of timing recovery systems. They can also be used to analyze the effects of impulse noise on the operation of satellites radio communication systems.

In the article, the problem of the organization of counteraction to ignorance of anti-radar missiles on a source of radiation of electromagnetic waves of a radio range is considered. Counteraction provides statement of active misinforming interference radiated from the point of space which is not coinciding with the position of protected radio station. Synthesis of the parameters influences of interference is made with use of differential game model. By results of definition of optimum strategy for antagonistic players: coalitions of the director of interference both protected system on the one hand and a self-guided missile on another are defined requirements to parameters of the interference maximizing the minimum prize of system of counteraction. It is shown, that such interference should be coherently connected with parameters of the signal radiated by protected system. Thus, the interference should have double frequency and peak manipulation which should provide periodic change of level of the received signal in the antenna placed on the head of the missile. In one half- cycle, the interference should exceed level of inadvertent radiation in side lobes of the radiation pattern of protected radio means. In other half-cycle interference level should be below radiation level in side lobes. The period of peak manipulation should be equal to the distinction of interference frequencies and the signal of protected radio means. In the actual conditions of the withdrawing simulating interference with the parameters defined on the basis of the differential game decision, the kinematic trajectory — which is aiming in essence — shows the analysis of modelling results of missile movement which leads to the miss.

The purpose of the work is to increase the simulation speed of decoding low-density parity-check codes and to reduce the memory requirements for storing check matrix and array of variable and check nodes. Simulation speed increases through the use of low-level language which is introduced into a Simulink model by means of S-function approach. In this work, C language is used. The gain compared with the MATLAB language is 20 times. Moreover, low-density parity-check matrix contains mostly zero elements which are not involved in the decoding but require a lot of memory. This paper proposes a method for storing positions of only nonzero elements and the number of units in each row. This measure has reduced memory consumption in 100 times; in addition to it, small internal arrays can increase the processing speed.
The described decoding algorithm and decoding results can be used in the curriculum. The developed code written in C is used in software implementation. The developed model is used to simulate different decoding processes. The designed technique allows one to reduce memory consumption. The speed of decoding corresponds to the similar block in the genuine Simulink library.

During the start of various purposes air vehicle it is necessary to determine their orientation within the small deflection angles (from a few to tenths of a degree). Radio devices do not provide the required accuracy of measurement and the use of satellite navigation systems (GLONASS, GPS) on board of certain types air vehicle is not always advisable. To solve this problem it is proposed to use optic-electronic systems (OES) designed for space technology, which include laser radar systems, laser beacons and cameras. Equipments of this type can be used in the field conditions and not only in the special test areas. A review of OES main models is performed. These models are used to determine the orientation of spacecrafts during rendezvous and docking, and their basic technical characteristics including precision measurement of the deviation angles are provided. The use of OES active type (laser radar systems) and the installation of optical response devices on the air vehicle (check points — corner reflectors or laser beacons) to
measure an angular method were substantiated. Expressions for the calculation of the angles of pitch, yaw and roll on two-dimensional images of check points and errors estimation were obtained. Comparison of errors for air vehicles deviation angles determination in the range between 10 and 300m shows that the accuracy of provided measurements coincides with existing analogues. The estimation of interference impact from air vehicles propulsion torches on useful signal power and OES range of measurements was carried out. Realization options for OES based on semiconductor laser diodes and photo-integrated matrix based on CCD or CMOS, and also with optic-mechanical scanning by narrow beam and with photo detector based on an avalanche photodiode was considered. The performed calculations show that the use of these technical solutions provides the range of OES measurements up to 300 m. Further assessment of OES capabilities will require consideration of factors determining the atmospheric trace state within measurements area.

Problems related to the movement of the aircraft or missile on the sliding bearings along the elastic rail beam arise at the present time in aviation engineering. For example, the launch of modern unmanned aircrafts in some cases is carried out by the rails.
The aim of this work was to research the process of aircraft lift-off from the beam and kinematic parameters determination. While considering the final stage of movement when the first bearing descents, it was taken into account that an extra grade of freedom appears at the expense of aircraft rotation towards the second bearing which hasnt left yet. The evaluation of accuracy of numerical algorithms and the estimation of convergence results were carried out.
In the process of problem solving there were considered transverse unsteady vibrations of elastic rail beam along which an absolutely rigid machine moves under the traction force. The forces acting on the beam from the bearings sides are found from the equations of motion taking into account the displacement of the beam. The displacements of the beams were presented by the Ritz method. The initial deflection of the beam is presented in the form of a springboard. There were obtained the equations of beam unsteady vibrations in generalized coordinates at all stages of the aircraft movement along the rail. The problem of machine kinematic parameters finding during the lift-off from the beam was solved for different variants of movement final stage.
The example of calculation is considered. The influence beam and machine of various parameters on the beam displacement, the transverse velocity and angular velocity of machine rotation when leaving the rail was investigated.
The result of the work is an estimation of beam and machine different parameters influence on aircraft lift- off ballistic parameters taking into account the particularity of movement final stage.

Most heated aircraft units are protected by thermal isolation. Isolation prevents structure from overheating. However, sometimes isolation could be locally destroyed. After that, discrete temperature field can appear in primary structure of the aircraft. It leads to a transient structure deformation and stresses. Also destroyed isolation has lower mass and it can affect the dynamic properties of aircraft thermal protection. All this factors are reasons for accident with space shuttle «Columbia». In fact, the thermal protection is covered on internal pipelines of aircraft. Pipelines isolation destruction leads to the unpredictable consequences. Therefore, dynamic behavior research on aircraft primary structure with local destroyed isolation has practical use.
Aircraft (or spacecraft) structure is simulated with the help of long thin-walled structure. Natural and enforced vibration analysis is provided for long thin- walled structure with local damaged isolation. Thermal protection coating has incomparably small stiffness characteristics in comparison to the supporting structure. Isolation destruction leads to the change of structure mass and inertial properties but it doesnt have influence on stiffness properties. This method of analysis also provides solution to hydroelasticity problem of fluid movement in conduit.
This problem is too comprehensive to be described in one article. Only one aspect of this problem is considered in present article. It is dynamic behavior of long thin-walled structure. This research consists of some parts.
Therefore, research on thin-walled structure dynamic behavior is illustrated by examples. The first example illustrates link between the size of isolation damaged zone and structure spectrum of eigenfrequencies. The second example illustrates forced oscillations under the influence of harmonic force. The third example demonstrates frequency response of thin- walled structure and the fourth contains research on the critical velocity of pipe fluid.

Purpose
The research into the fundamental astrometry problem of high-accurate interpolation and prognosis of the rotational-oscillatory motion of the deformable Earth according to the International Earth Rotation and Reference Systems Service (IERS) observations for navigation applications.
Design/methodology/approach
The improvement of the ephemeris-time support accuracy problem in satellite navigation systems is connected with the fundamental problem of the Earth orientation parameters (EOP) determination: with the Earth pole oscillations and length of the day variations in the short time intervals in particular. In this paper the mathematical models for the Earth pole coordinates shift and for the time scales desynchronization between UT1 and UTC (dUT1) are shown, that agree with the IERS. The significant improvement of the prognosis accuracy is achieved by taking into account the perturbations of the oscillatory-rotational motion of the Earth in the space vehicle (SV) motion equations.
Findings
The graphics of the diurnal prognosis error of the Global Navigation System (GLONASS) navigation satellites are presented that is caused by the Earth pole shift. A comparative analysis shows that the perturbations of the navigation satellites orbits are comparable to the gravitational impact from the Moon and the Sun and to the perturbations from the irregular part of the Earth gravitational field.
Research limitations/implications
High-accuracy information about EOP is required on the SV for solving navigation problems. It is shown that the introduced models are able to provide self- sufficiency for the SV to calculate EOP on its own.

Influence of high-frequency vibrations on stability of stationary motions of mechanisms used in aviation and cosmic techniques is a very actual problem. Vibrations can change the nature of stability and cause new dynamic effects. Furthermore, friction should be taken into account when studying real systems.
We consider here a model problem of a motion of a rigid body, one point of which (a suspension point) accomplishes vertical harmonic high-frequency oscillations of small amplitude. A force of viscous friction is assumed to act on the body, with a torque proportional to its angular velocity. Stability of two relative equilibriums for which the center of mass and the suspension point lie on the same vertical is studied. The investigation is carried out in the framework of the approximate autonomous system of differential equations of motion obtained in the paper [11]. Conditions for stability of the vertical relative equilibriums in the case of absence of friction have been also obtained in [11].
Three special cases of the body mass geometry are considered here when the bodys center of mass lie on one of the principal axes of inertia or in the principal plane of inertia, or the body is dynamically symmetric.
A characteristic equation of linearized equations of a perturbed motion has a zero root. The other roots have negative real parts if friction is small and the vibration frequency exceeds the special value (for the upper equilibrium) and has any value (for the lower one). The nonlinear parts of the whole perturbed equations are proved to have a special structure. This is so-called special case of one zero root in terminology of [1]. Investigation of stability is carried out by using algorithm of this monograph. It is shown that the equilibriums under consideration are asymptotically stable in five variables and stable (not asymptotically) in the sixth (critical) variable.

On the base of the Love model and the classical Fourier thermoconductivity theory the analysis of the thermoelastic behaviour of the geometrically irregular plate is carried out. Two opposite edges of the plate are exposed to quick change in temperature fields and force stress on space coordinate. Such temperature fields and force stress are described as multiplication of suitable functions by the Heaviside function. These functions at force stress points, when the Heaviside function is not defined, become zero along with their derivatives until the required order.
The temperature field, in conditions of convective heat exchange via surfaces and boundaries exposed to quick change in temperature, is defined in the plate beforehand. One temperature jump over one of the surface caused by partial destruction of thermal shield is supposed. The solution of the boundary problem of thermoconductivity has been obtained in a closed form suitable for quantitative analysis and sequential solution of incoherent thermoelasticity.
The solution of the thermoelacticity problem is reduced to integration of inhomogeneous irregular differential equation with variable coefficients in a form of the generalized Dirac function and its derivatives. By the method of the sum single trigonometric series with variable coefficients and different functions for satisfying boundary conditions the solutions are reduced to the integration of inhomogeneous ordinary differential equation with irregular coefficients. The fundamental system functions are obtained for an irregular equation. The solution is written in a closed form. The conditions of fixing two edges of thermoelastic system at rest can be any of the known in the theory of the plate.
On the basis of exact solutions the 3D figures of thermal surfaces and bending surfaces are obtained for different values of physical, mechanical and geometrical parameters. The configuration of surfaces greatly depends on the directions of quickly changing forces and a temperature stress.
The regions of largest bending coincide with the regions of largest temperature difference in thickness. The bends are less sensitive to the intensity forces stress on the boundaries. It is worthwhile noting that practically in all considered cases the increase of parameter of the relative height of the confirmatory rib decreases the value of bending if the heat insulation of the surfaces is not broken.
In the paper the given solutions of the incoherent thermoelasticity of the geometrical irregular plates exposed to the quickly changing loading are aimed to the analysis of thermoelasticity behaviour of the thick constructions required in the various technological areas.

Subject of study in the article are models and methods of analysis and evaluation of functional parameters of successive line production process. The aim is to assess the technological feasibility of the planned production processes through the analysis and consideration of the characteristics of individual process steps and equipment used.
Tasks it is been: ground of expedience of application of mass service theory for the mathematical design of technological processes; development of model of the queuing system, describing successive technological processes for the receipt of functional and cost descriptions of production; determination of mathematical expressions for the receipt of cost descriptions of production processes with the purpose of determination of equipment composition for providing of the set productivity and necessary amount of the same type equipment.
In-use methods are a mathematical design on the basis of theory of mass service, in particular are probabilistic models of the multiphase queuing system.
Following results were obtained. It is shown that, when assessing the feasibility of the planned production of the manufacturing process it is advisable to use mathematical modeling, since it is possible to determine the qualitative and quantitative characteristics of equipment and tools, based on which the analysis is made possible TP and rating them on the specified criteria, such as cost, time, volume release and other. The mathematical expressions for the main characteristics that define the state of the production system as a multiphase queuing system are obtained. The integral characteristics of production, such as the average &amount of work in progress and the average cycle time are determined. The cost queuing model aimed at determining the optimal level of functioning of the production system is given. The steps of method for determining cost characteristics of processes are proposed on the based on the above models.
Conclusion. Thus, we obtain the possibility to simulate the stream processes of modern production and evaluate the time and cost parameters for the feasibility analysis of requirements in the development plans of production. The results can be used in the development of information technology and automation tools for the planning and control of production processes.

Purposes
1. Analysis of restructuring strategy realization of the aviation industry as a whole especially in helicopter
industry
2. Prove working hypothesis that the creation and development of strategic competence centers will improve the competitiveness of aviation corporations, especially under conditions of uncertainty and instability that exist in todays global economy. Identification of unique competencies enterprises comprising a holding may become one of the levers to enhance the functioning both the industrial enterprise and the entire holding
Design/methodology/ approach
Strategic management methodology is used in a part of mission formation, strategy perspective. Opportunities of development strategy aviation industry are considered on the example of holding ОАО «Helicopters of Russia». Influence creation competences centers are certain on the increase of forecasting efficiency for "Stupinsk"s machine-building productive enterprise" included in holding. Degrees and interdependence of restructuring in the aviation industry and creation of business units which coordinate activity of the branch enterprises and the competences centers, defining the contribution of each enterprise. The usage of business models for parameter formation of the competences centers in the machine-building enterprise of the aviation industry are shown.
Originality/value
The necessity of creation centers of competence in engineering aviation industry is shown in article. A hypothesis was formed about the influence of competence centers on the efficiency of industrial enterprises, Business model was developed to determine the contribution of machine building company in the aviation industry job holding, Tasks for the competence centers were formed. Mission and strategic vision , given the forecast efficiency of the implementation the program of strategic change within the competence center " units " of the enterprise.

The paper presents the analysis of the available power consumption by the fenestron in different flight modes as well as the analysis of the change of the single-rotor helicopter weight with the installation of the fenestron. The analysis aims at determining the values of the design parameters of the fenestron and evaluating the rationality of its use on helicopters of various weight categories.
The parameters of the fenestron, which require value selection, include the shroud duct parameters (parameters of the duct behind the fenestron rotor) and the fenestron rotor parameters. The weight of the cargo, which is transported by the helicopter with a fenestron over a given range, is used as the indicator of its effectiveness. It is also used as a criterion for the selection of the fenestron parameter values. The weight of the cargo, which is transported by the helicopter with a fenestron over a given range, is determined from the aircraft existence equation

Here  is the helicopter take-off weight;  the helicopter empty weight (it is mainly determined by the helicopter structure weight); is the weight of the fuel, which is required to carry out the flight; is the weight of the operational load
The fenestron parameters influence the helicopter engines power and thus the weight of fuel, which is required to carry out the flight, accordingly. In its turn the engines power has a substantial impact on the weight of the helicopter structure as a whole. Therefore it also influences the weight of the cargo, which the helicopter can transport.
The optimal value of the ratio between the fenestron rotor diameter and the main rotor diameter was determined as a result of the parametric studies. It was pointed out that the increase of the main rotor disc loading pнв corresponds to higher optimal values of . These values are equal to 0,07..0,08 for light helicopters and up to 0,09...0,11 for helicopters, which belong to the medium weight category.
The relative diffuser length , which is calculated in fractions of the fenestron rotor radius , proves itself in a similar way in the helicopter characteristics. On one hand, the increase of the relative diffuser length leads to the decrease of the power that the fenestron consumes in hover mode. On the other hand its increase results in the growth of the lateral fenestron dimension and thus leads to the rise of the helicopter parasitic drag in horizontal flight. It was established that the optimal relative length of the diffuser depends significantly on the requirements concerning the design flight range of the considered helicopter. The bigger is the design flight range the shorter the diffuser should be. Thus, for example, it is possible to make a relatively long diffuser for a light helicopter with the relatively short flight ranges (L = 200...400 km). Flying over longer ranges requires the decrease of  value down to 0,4...0,5.
The paper also presents the comparative analysis of the characteristics of a helicopter with a fenestron and a helicopter with a conventional tail rotor. The results of this analysis were used to obtain the dependencies, which allow the designer to evaluate the rationality of installation of the fenestron on helicopters of various weight categories. Thus, for example, the use of the fenestrons on light helicopters almost does not affect the ratios of their payload and fuel weight to take-off weight. The losses (decrease) of the ratio of the payload and fuel weight to take-off weight can amount up to 10% for helicopters of medium weight category with maximum take-off weight of 10000 kg and design flight range of 600 km. The use of fenestron is not rational for helicopters with bigger take-off weights due to even bigger losses in their ratios of the payload and fuel weight to take-off weight.

The paper reviews a possible solution to the problem of power supply of the actuators of the main flight control surfaces, which ensures the safe controllable flight of a passenger airplane in case of failure of all of its propulsion engines. In this critical emergency situation the performance of the main sources of the hydraulic power degrades and the flow of the hydraulic fluid, which is generated by the pumps, approaches zero. If the propulsion engines can enter the autorotation mode, the process of degradation of hydraulic power to supply hydraulic power to the actuators of the main flight control surfaces. Such power supply could be combined with the power of the emergency ram-air turbine, which transforms the ram air energy into the hydraulic power by means of a turbo-pump unit.
The paper analyzes the process of degradation of the hydraulic power supply to the flight control actuators. This analysis shows that there are some time intervals, during which it becomes impossible to control the flight. The paper suggests carrying out a modification of a typical composition and configuration of the airplane onboard hydraulic system, which would ensure the controllability of the flight. The paper adduces the results of simulation of operation of the flight control actuators in case of all propulsion engines failure. The simulation is carried out for a typical composition and configuration of the onboard hydraulic system and for a modified system, which incorporates an additional emergency pump with a direct current electric drive. This additional hydraulic power source is only engaged temporarily during the time when the ram-air turbine and the turbo- pump unit are extended into the air flow. The additional hydraulic power source is disengaged after the feed rate of its pump stabilizes. The paper reviews a model, which simulates the process of degradation of the hydraulic power supply. It also adduces the results of simulation of actuation of the flight control surfaces, which was carried out with the help of the aforementioned model. These results show that the operation of the flight control actuators, which are driven by the harmonic input control signals, may be distorted substantially in case of the propulsion engines failure and RAT engagement. The abovementioned distortions of the processes of actuation of the flight control surfaces disappear after the incorporation of the emergency power unit into the onboard hydraulic system of the airplane.

Scope of research
The attitude control system of angular motion of the launch vehicle (LV) is the subject of study in this work. The problem of the statistical analysis of a vehicle angular motion of the LV taking into account a horizontal wind as a correlated in altitude random function is solved in the work. The second order shaping filter is used for the wind description as a correlated random function of the vehicle flight altitude [2]. Statistical characteristics of this process have to be taken into account in the analysis of normal accelerations experienced by the LV during its flight in windy atmosphere, both at the structural design of a new LV or during modification of existing LV.
Research topic
Technique of the statistical analysis of 1st stage LV motion taking into account random wind loads.
Purpose
Purpose is to develop technique of the statistical analysis of controlled motion of the 1st stage LV in presence disturbances due to horizontal wind in an atmosphere with the use of moments equation method.
Methodology
The technique provides calculation of statistical performance indicators of stabilization system by use the moments equation method. To imply the method it is nessecary to obtain the differentian equations for the augmented state vector describing angular motion of LV in Cauchys normal form. Augmented state vector also includes two variables of the second-order shaping filter describing the wind as a random process. The mean vector and covariance matrix of the augmented state vector are estimateg by integration of quazi-linear ordinary differential equations for these characteristics.
Statistical characteristics of all state variables of an attitude control system, as well as characteristics of output of interest and variance range, are results of calculations. The proposed technique is implemented in a calculation program development in MATLAB/ Simulink environment.
Results
Statistical characteristics of output of interest -normal acceleration and all component of a state vector of LV are obtained as the results of calculations.
The results obtained with the use of proposed techniques are compared with the similar results calculated by the Monte-Carlo simulation and the approximate «enveloping» methods.
It is shown that the methods of equation moments and Monte-Carlo give similar results, although the Monte-Carlo method is significantly more computing time-consuming. Easy for implementation method of «envelops» can lead to receiving the underestimated estimates of possible values of normal accelerations of LV.
Research implications
Relevance of the work is determined by the dynamic development of space research, which require creation of various satellites, which differ both in weight- dimension characteristic and design-layout scheme. In addition, a number of space research projects requires a simultaneos insertion into orbits of several spacecrafts that places specific requirements to design of head part of LV, in particular, leads to the use of large head fairing (nose cone). It is known that the normal g-force experienced by LV increases with the increasing of fairing volume of a nose cone. In that case it is necessary to take into account all the disturbing factors as accurately as possible, and the major disturbing factor for the 1st stage of LV is the wind.

Simulation is well known and highly used instrument for majority of space mission control tasks. Special software tools are used for development of spacecraft simulation models. These packages consist of a number of simple elements and algorithms designed for such elements synthesis. But some disadvantages were found in existent tools. The article is devoted to the new simulation models development method without such disadvantages. This method is based on structural decomposition into the simple models with one or a few short equations.
Presented in the article method uses three-stage structural model decomposition. Spacecraft dividing circuit into elements level is used as a scheme for first stage, each model element is separated to the simple models (SM) at second stage, and at the last stage each SM is divided to set of fragments. SM is a model of element details (such as iron bolt or electric cable) and SM fragment describes SM participation in concrete physical domain (e.g. electrical or mechanical). Represented decomposition scheme provides three radically different viewpoints: on each element of dividing circuit, on entire spacecraft scheme of concrete physical domain and on cross-domain interactions. Pairs of SMs could be linked with bonds. Each bond transfers parameters values between connected SMs.
Element model contains a number of interfaced and internal SMs. Interfaced SMs could be connected with other models of SMs elements and internals describe elements behavior. Models of elements of the same dividing circuit level could be combined into a model of spacecraft subsystem or system. Equations of different models could be computed on different processor units and cooperated with bonds mechanism.
Each SM with fragments contains one or few short equations of specified physical domains and cross- domain interactions. Models of elements could also contain the short particular equations to specify its behavior. Generally, such equations are logic, algebraically non-linear or have the first order differential as it mentioned in the article. Set of equations could be numerically solved with well-known math methods.
The article describes the process of abstract spacecraft propulsion system model creation. Definition of model parts including fragments, significant parameters, formulas and bonds is exposed. Examples of code are given. Propulsion system specific processes simulation graphs are shown.
With presented method it is possible to develop modern software tool for subsystem and entire spacecraft simulation. Models under creation will be tunable, inter- operable, scalable and re-usable. Such models will be useful for spacecraft mission control tasks and, especially, for integrated model-based health management.

Subject
Interplanetary trajectory of the spacecraft
Title
Trajectory optimization of the spacecraft with electric propulsion to the Jupiter flight using the sequence of two gravity assists near the Earth
Purpose
Trajectory design analysis of the Jupiter spacecraft path with electric propulsion using the sequence of two gravity assists near the Earth and comparison of this path with simpler ones. Development of optimization method for interplanetary flight trajectories with the help of gravity assists sequence.
Design/methodology/approach
Basis methodical approach is the Pontryagins maximum principle. Formulation of four point boundary value problem with consideration of gravity assists optimality conditions performing.
Continuation method is the basic idea for the solution to boundary value problem.
Originality/value
Optimality conditions for both possible cases of gravity assist are obtained and performed (the case where the gravity assist is performed with the minimum perigee altitude of hyperbole flight path and the case where this altitude is higher than the minimum perigee altitude).
Formulation of four point boundary value problem including the optimality conditions for gravity assists.
Results
Four point boundary value problem is presented. Optimal conditions for gravity assists are presented.
The spacecraft trajectory characteristics to the Jupiter flight using the sequence of two gravity assists near the
Earth are shown.
Trajectory design analysis to the Jupiter flight with for transport system which is based on the launcher "Angara — 5′, chemical upper stage KVTK and electric propulsion using the sequence of two gravity assists near the Earth is presented.
Comparison of trajectory design analysis for three flight paths (direct flight, one gravity and two gravity assists near the Earth) to the Jupiter flight is presented.
Practical implications
During the development of programs for advanced space studies, for research of the giant planets. In the analysis of complicated interplanetary flight paths including gravity maneuvers.
Conclusion
The method of trajectory optimization of the spacecraft with electric propulsion to Jupiter flight using the sequence of two gravity assists near the Earth is developed. Using the necessary optimal conditions of maximum principle the formulation of four point boundary value problem is presented.
The peculiarity of the formulation is consideration of two feasible gravity assists realization (the case where the gravity assist is performed with the minimum perigee altitude of hyperbole flight path and the case where this altitude is higher than the minimum perigee altitude). It is shown that flight to Jupiter using the sequence of two gravity assists near the Earth can significantly increase the mass of spacecraft delivered in the vicinity of the Jupiter about 560 kg (from 6425 to 6985 kg) in comparison with the flight path using a gravity assist near the Earth.

The resonant electric steering drive (RESD) can be applied as a part of control systems of all classes of rockets rotating on a list. Existing samples of RESD still far from perfect and their weight and dimensions are redundant. The purpose of this work is receiving a quantitative assessment of weight surplus of the best sample of RESD known now. In the course of work the problem of receiving a quantitative assessment of a stock of power and weight of the executive engine (EE) is solved. EE is the basic element of the drive defining its specific characteristics. Research is conducted by method of modeling of an operating cycle of the hypothetical guided missile. In this paper the questions of creation of the RESD imitating model which takes into account the thermal processes, occurring in EE and its research are considered. The model is created in the environment of Simulink entering into a mathematical MATLAB package. Models of entrance influences of the drive changing in time and loadings of its output link are constructed. The power stock of the EE was estimated at temperature of an overheat of windings and on energy of losses. The simulation results showed that the power of the EE in more than 10 times greater than those required output. It confirms that samples of RESD existing now possess the big excess weight and dimensions which can be significantly reduced at the expense of decrease in weight and dimensions of EE. The obtained results allow to carry out works in the direction of definition of extremely achievable specific characteristics of the steering drive at the present level of technology.

This paper is about the method of turbo machinery blades dynamic optimization. The method allows receiving the necessary dynamic characteristics of the blades taking into account long-term static strength and requirements of gas dynamics and structural constraints.
For optimization the special software is developed using programming language APDL in ANSYS. The software automates all modification processes of the original blades structure and includes a procedure for the automatic transfer of quasi-3D model cross-sections point coordinates to the 3D modal analysis ANSYS package and back. The process of blades dynamic optimization is performed using software LMS OPTIMUS. The optimization process is completed when all requirements to the structure are satisfied with a given accuracy.
The software allows performing the dynamic optimization of the blades as part of wheel taking into account shroud, interlock and disk. As the variable parameters of a blade airfoil, chords, maximum thickness, global coordinates of the mass centers and setting angles of the blade cross sections are used. Cross sections point coordinates are obtained for each of optimization iteration according to varying parameter. Two automated methods of the blade dynamic optimization are proposed. The first way assumes calculation of influence factors determining the section parameters change against the frequency (or several frequencies) change. Basing on the analysis of the influence factors their optimal combination is obtained. This approach takes a little time (a few hours). The second way is to use the standard methods of multi-criteria optimization, realized in the package LMS OPTIMUS. Therefore, the time of the optimal design obtaining significantly increases (several days) but the quality of the optimal design becomes better. The example of typical blades dynamic optimization using both approaches is performed. The original blade has a small margin (<2.8%) in relation to the possible resonance of the first oscillation mode with the harmonic k = 4 to the rotation frequency. By optimization using the first method the margin increases to 10% and the blade mass grows by 8.2%. The second optimization method leads to the margin increase to 9.8% but the mass grows only by 4%. The use of the proposed methods and modern software packages for multi-criterion optimization allows to automate a process of the blades dynamic optimization, reduce the time and laboriousness of the design process, and improve the quality of the designed wheel units.

Purpose of the work is development of mathematical model (MM) of calculation of thrust-economic and weight-dimension characteristics of power plant (PP) for the unmanned aerial vehicle (UAV) with a turbojet.
Methods of the work are the analysis and numerical experiment.
Results of the work
At present, development of UAV is one of the priority directions of developing of aviation in the world. High-altitude UAV with long-range and flight duration are especially required for Russia, taking into account the extent of its territory.
Authors developed dual-flow turbojet engine mathematical model for calculation of thrust-economic and weight-dimension characteristics. This MM enables carry out calculations with high precision for minimum time. The MM possesses high stability for carrying out optimizing researches.
Distinctive feature of developed MM is its multi-schemes. Performances of turbojet and dual-flow turbojet (with stream mixing and without it) with various quantity of rotor shaft of the engine (single-shaft, twin-shaft and three-shaft schemes), with afterburner and without it and also dual-flow turbojet with add stage and gear box can be calculated using this MM.
Test calculations of performance of serial engines (D-30KP and PS-90A) which showed high adequacy and reliability of this MM were carried out with using developed the MM of PP.
Parametric investigations of basic variant of dual-flow turbojet were performed.
Application area of results
Results of this work can be used in the scientific and design organizations engaging in development of draft proposal for advanced UAV, in the ordering Air Force organizations and the industry at justification of requirements for new samples of the aircraft equipment, and also in aviation institutes of higher education at improvement of educational process.
Conclusions
As a result of the conducted researches it is possible to draw the following conclusions:
1. Development of aviation engines of various types for domestic UAV, which satisfy modern requirements for specific fuel consumption, specific weight, etc., is actual.
2. For the solving a task on formation of preliminary technical shape of power plant of the UAV it is developed MM, making it possible to solve a wide range of engineering problem at calculation of aviation engines of various schemes.
3. The carried-out verification of developed MM showed its good adequacy and reliability by comparison of design and passport data of a number of serial aviation engines.
4. The performed parametric researches showed that on a number of work process-related parameters of dual-flow turbojet there is their ambiguous influence on thrust and specific fuel consumption. Therefore carrying out of calculating researches on formation of optimal technical shape of PP on the basis of dual-flow turbojet in UAV system is supposed in what follows.

Results of the plasma potential, temperature of electrons, radial ion flow density and ion energy experimental study in the vicinity of the stationary plasma thruster (SPT) model exit plane are represented in this paper. This SPT model with external accelerating channel diameter of 85mm was developed at RIAME MAI and operated with discharge voltages of (300-800)V and mass flow rates through the mentioned channel of 2.1mg/s and 2.5 mg/s of Xenon. Plasma parameters were measured by several flat electrostatic probes positioned in the vicinity of the thruster exit plane along different radial directions and at the same distance fr om the thruster axes corresponding to the typical distance of cathode position from the mentioned axes. Energy of ions in the vicinity of thruster exit plane was measured by the retarding potential analyzer (RPA). It is shown that character of the plasma parameter dependence on the discharge voltage changed significantly with the increase of the mass flow rate through the thruster accelerating channel from 2.1 mg/s to 2.5 mg/s and that the mean energy of ions was around (80-120) eV and weakly dependant on the discharge voltage within the range of the discharge voltages of (300-800) V. Thus, energy of ions in radial ion flows moving in the vicinity of the thruster exit plane is sufficient to sputter structural elements such as cathode.
There were considered the possible mechanisms of ion generation with relatively high energy discovered in the radial ion flows and it is shown that elastic scattering of the accelerated ions on neutrals can not provide the measured level of the ion current density. Estimations show also that the acceleration of ions appeared near the exit of the accelerating channel by the potential drop in plasma volume due to the electron pressure gradient can not result in the measured energy of ions. Therefore it is concluded that the main source of such ions could be an exit part of the accelerating layer positioned outside the accelerating channel of the studied SPT model wh ere the lines with constant potential near the boundary of the main flow can have configuration causing an acceleration in radial directions of ions appeared near the mentioned boundary.

With use of special samples simulators pilot researches of growth rate of a crack in the conditions of action of the residual squeezing tension which size is characteristic for the high-loaded details of aviation engines are conducted. On the basis of the settlement pilot studies including definition on models of high level of intense of the deformed condition of special samples simulators with cracks and definition, by means of a fractography, a step between grooves, the kinetic charts of growth rate of a crack in a material of special samples in the conditions of action of residual compression stresses tension are received.
Values of the stress intensity factor in points of the front of a crack were determined by results of elastic calculation and were corrected, for taking note of residual compression stresses. Results of experimental determination of growth rate the cracks received on a basis the fractography researches of real disks, in the conditions of action of residual compression stresses practically coincides with results of the calculations executed with use of the offered way that confirms possibility of settlement forecasting of durability taking into account the offered way.
The way is developed for taking into account influence the residual compression stresses arising in zones of concentration of tension as a result of hardening of a material at its elastic-plastic deformation on the growth rate of a crack
The held testing showed good coincidence of results of the calculations executed with use of the offered way, to the data obtained on the basis of pilot studies of growth rate of cracks in disks of aviation engines.
The received results testify to influence of residual compression stresses on the growth rate of a crack and to need of the accounting of residual compression stresses for zones of concentration of tension at an assessment of durability of critical details of aviation engines.

Modern and perspective space vehicles (SV) widely use electric rocket engines (ERE) for an attitude control and space vehicle orbit correction as well as in the capacity of sustainers. As it has shown probing, under certain conditions data ERE except thermal noise radiation can form radiation in the form of a random pulse sequence in microwave band.
The considerable quantity of methods of struggle against an impulse noise is known. The most known algorithm is the base algorithm in which the received signal after a filtering in a wide band of the receiver is subjected to a limiting, and then filtering in the band matched with a spectrum of an information signal (i.e. in the matched filter). The given algorithm of struggle against an impulse noise has shown high enough efficiency and is widely used in practice in a case when the width of a pulse noise spectrum is much more than width of an information signal spectrum.
The offered algorithm differs fr om the known one which detection and impulse noise cut implements for. Signal interpolation on the interval wh ere an impulse noise is besides made. The algorithm singularity is its implementation in the form of the interconnected quadrature channels.
Taking into account nonlinear character of an impulse noise compensation algorithms use of analytical methods for probing of their characteristics is coupled to significant mathematical difficulties. In the present paper on the basis of simulation methods the comparative analysis of efficiency of several pulse noise compensation algorithms is carried out and questions of their parameters optimization are observed.
Mathematical models of a communication system have been realized in the environment of MATLAB/ Simulink. Results on an instance the channel of digital communication the Earth-SV are observed. It was supposed, that modulation in the form of the FBPSK is used by filters of type of the raised cosine.
The carried out analysis and results of modeling show, that the offered algorithm of the impulse noise quadrature compensation in the conditions of combined affecting AWGN and pulse noise allows to gain BER (Bit Error Rate) in comparison with the base algorithm. So, for the optimum thresholds of limiting and detection pulse noise at a Eb/No on a receiver input 5 dB the benefit will be approximately twice, and at a Eb/No 10 dB — to 4 times and over (for signal to pulse noise ratio from 0 to 5 dB).
Observed compensation algorithm is expedient for using in a combination to the channel of measurement of current level of the AWGN and an impulse noise for detection of optimum value of a threshold. The received results can be used at designing and selection of parameters of space communication systems.

The aim of work was to evaluate CST I- and CST А-Solver module accuracy for calculating reflection cross-section (RCS) of objects with typical geometric form. The methodology of accuracy evaluating comprised: experimental measuring of RCS angular diagrams for work objects, diagrams calculating in CST I- and CST А-Solver, calculating of deviations between calculated and experimental diagrams per ranged characteristic coefficients, root-mean-square errors, average relative errors. Ranged characteristic coefficient characterized linear dependence degree between compared diagrams and ensured root-mean-square error and average relative error to be good evaluation of CST accuracy. In the work, it was supposed, that errors are good evaluations, if ranged characteristic coefficient value is equal or bigger, than 0.7. The work was executed for electromagnetic wave (EMW) with following frequencies: 2, 3, 4, 8, 10 GHz; both horizontal and vertical EMW polarization was considered to make the investigation complete.
Cylinder and corner reflector were work objects with experimental data measured in ITAE of Russian Academy of Sciences. So it was shown, that CST I- Solver and CST А-Solver calculate RCS for objects of typical geometric form with appropriate accuracy, which is conformed to accuracy of modern experiments. Difference between CST I-Solver accuracy and CST А- Solver accuracy was non-significant — at a level of 0,6dB. Since CST А-Solver demands of less time for calculating, than CST I-Solver, CST А-Solver is preferred to evaluate RCS for object of typical geometric form quickly and sufficiently exactly. CST I-Solver is preferred being demanded sufficiently exact evaluation of object RCS diagram character.

This article describes the problem and the current state of the transceiver modules (TM) of active phased array antenna and digital antenna arrays (DAA), as the next stage of their development. Structure of the DAA in which TM consist of highly stable grid frequency synthesizers synchronized by digital reference generator was considered. Exception complex microwave wiring and traditional phase shifters change on digital devices beamforming (quadrature modulators) requires substantiation requirements for hardware components. Particular attention is paid to the frequency range from 4 to 12 GHz.
The TM is the key element in proposed DAA and its characteristics depend on power amplifiers. Implementation of multibeam increases the potential energy required for DAA, which leads to an increase in the required TM output power of 2050 watts. Accurate calibration of the DAR and the creation of multibeam patterns on the transfer requires a small discrete value of the amplitude and phase, which imposes restrictions on the operation of the power amplifierset out the main power amplifiers requirements for on- board DAR — frequency range, linearity, output power level.
The article describes GaN transistors, which usage requires new development and modeling methods for designing monolithic integrated circuit of power amplifiers; the advantages and disadvantages of proposed TM for DAA are discussed. Features of the use of multisection transistors and assembly technology for power GaN transistors are discussed. High cutoff frequency of GaN transistors allows creation of ultra- wideband power amplifiers up to 20 GHz and its efficiency up to 30%.
High values of power added efficiency and power density per unit gate width using AlGaN / GaN pHEMT provide the significant reduction in weight and size of power amplifiers based on them that determines their applications for on-board radioelectronic systems. Keywords: GaN power amplifier, digital antenna array, digital beamforming.

Theoretical and experimental approach for heat chamber parameter choosing for thermal tests of active phased antenna array digital radar module operating on nonpressurized Earth satellite is discussed. The suggested approach includes three theoretical and experimental steps. At the first step the digital radar module thermal behavior is tested in heat vacuum chamber that most closely simulates the environmental conditions of the nonpressurized Earth satellite. At the second step mathematical simulation of the digital radar module heat processes in heat chamber is carried out. At the third step the algorithm is developed and the heat chamber parameters are chosen providing the digital radar module temperature field in the heat chamber identical to the temperature field of the same module in the heat vacuum chamber. The criteria of the heat and heat vacuum chambers thermal conditions identity is the criteria of minimum mean square error of the digital radar module temperature fields in the both chambers under the heat release sequence diagrams given in Technical Requirements Specification.
The study conducted proves that the chosen heat chamber is suitable for complex thermal tests of parameters of the active phased antenna array digital radar module intended for operation on nonpressurized Earth satellite. The presented approach opens opportunities to carry out the performed tests for various onboard radio-electronic equipment operating in spacecraft nonpressurized modules in heat chambers instead expensive heat vacuum chambers.

The paper considers determination of an aircraft position using the processing of carrier phase satellite radio navigation signals from several GPS/GLONASS (below referred to as global navigation satellite systems — GNSS) antennas mounted on the fuselage. The specific point of carrier phase measurements is phase ambiguity which makes it impossible to obtain attitude solution estimate using the measurements only at one instant of time. So, measurements from a certain interval of time are used to make the estimate consistent. The desired attitude parameters and phase ambiguity become observable due to the motion of the GNSS satellites constellation. The accuracy of attitude determination naturally depends on the accuracy of phase ambiguity resolution, which in turn severely depends on the constellation kinematics. These relations are under the study. The paper gives an approach to a priori accuracy of phase ambiguity resolution.
In addition to dilution of precision factors (DOPS) which characterize the position of satellite in view constellation at one instant of time and are well known in GPS navigation, here we introduce analogous integral characterization of constellation motion. These new quantities are called DOPCK (dilution of precision due to constellation kinematics). Their connection to the real accuracy of phase ambiguity resolution is shown. Quantities are introduced as follows

where tr() is the trace of the matrix , is estimated error covariance of ambiguity resolution, is a reference noise level and n is number of measurements. Positional dilution of precision factor PDOPCK is much like the same except for positional covariance matrix is used in formula. The higher the value — the less accurate estimate is produced, and vice versa. Processing of a survey data shows that every time phase ambiguity has an outlier, there is a leap in ADOPCK. This fact indicates that there is a real connection between new introduced quantities and phase ambiguity estimation.

Fig. 1. The temperature distribution in the DW (). Models: 1,2 «equilibrium»; 3, 4 «Nonequilibrium reversible»; 5, 6 «Nonequilibrium irreversible»; 1, 3, 5 initial droplet diameter =10  microns, 2, 4, 6 = 50 microns

а)

 б)

Fig. 2. a) The change of temperature in DW (= 298,15 К), b) The temperature in the BW (= 1000 К) on the phase VT plane ( = 101325 Pa, = 50 microns): . Model: × «equilibrium»; «nonequilibrium reversible»; solid line «nonequilibrium irreversible» (ДА equilibrium adiabat, УА shock adiabat)
Field of application of the results and conclusionsDeveloped computational algorithms and computer programs can be used for designing advanced propulsion ystems.

This article discusses the current state of the regional jet aircraft world market in terms of the major players customer base regional analysis — «Embraer» and «Bombardier», as well as the prospects of domestic products in this class — SSJ- 100 aircraft.
The authors came to the following conclusions in their paper:
1. Within the framework of the global transformation processes of regional jet aircraft market, the analysis of market behavior of the «Embraer» and «Bombardier» companies in its geographical aspect in order to assess its impact on the prospects of advancing of the domestic products produced by JSC «UAC» to the regional markets is becoming particularly urgent.
2. The customer base of the aircraft manufacturers can serve as an object of the investigation, characterizing the market behavior of the «Embraer» and «Bombardier» companies, i.e. the structure of the companies purchasing the aircrafts manufactured by them as a defined organization system, formed on the basis of the action of the objective laws of the market.
The analysis of the client base of the airlines exploiting"Embraer" ERJ 170/175/190/195 and «Bombardier» CRJ 700/900/1000 airplanes shows the substantial difference ( about 1.5 times judging on the amount) between"Embraer" and «Bombardier » according to the degree of the country coverage of the global regional jet-aircraft market.
At the same time, the distribution of airlines exploiting «Embraer» and «Bombardier» aircrafts also has significant differences within their customer base between each of these aircraft manufacturers. Client bases of the airline operators of the ERJ 170/175/190/ 195 and CRJ 700/900/1000 aircrafts do not intersect and there is only one airline («Lufthansa Cityline»), which at the same time includes ERJ190, ERJ195 and CRJ700, CRJ900 in its aircraft fleet.
The main (key) customers of the «Bombardier» CRJ700/900/1000 aircraft are 9 airlines, which account for two-thirds of these aircrafts in the world. The aggregate share of the «Bombardier» aircrafts, held in U.S. and European airlines is about 88 % of the existing fleet of these aircraft, therefore the market prospects of the Canadian company are largely determined by the situation on the European and U.S domestic markets.
In relation to «Bombardier», «Embraer» has a more balanced market presence in regions of the world. The demand and situation on the domestic markets of such countries as Canada, Brazil, China, Great Britain, Germany, Australia and the countries of the Persian Gulf will play a significant role in the «Embraers» market prospects.
3. The study of the foreign manufacturers client base allows to offer «thinner» approaches to the positioning of the Russian SSJ- 100 aircraft on the world market, regarding it within its comparison with the market opportunities of competing products -«Embraer» ERJ 190/195 and «Bombardier» CRJ 900/1000 aircrafts. At the same time the study confirms the correctness of the strategy of the SSJ- 100 aircraft manufacturer CJSC" Sukhoi Civil Aircrafts" about the importance and prioritizing of the markets of Asian-Pacific regions and Indochina.
It is shown that the usage of the traditional approach is not absolutely correct as CRJ 900/1000 aircraft «a priori» are the competitors of SSJ- 100, moreover it is not true for a number of regional and country markets. This assumption allowed the consideration of a new price positioning of SSJ-100 in relation to the CRJ 900/1000 on the markets of countries, where this aircraft is not presented.

The problem of selection of fenestron design parameters consists in finding such a combination of geometrical parameters of the fan, inlet and diffuser within the given overall dimension of the fenestron, which would provide the least possible power consumption of its fan at the given thrust value. The varied design parameters (parameters, which act as design variables) include the following: the fan radius R, the relative curvature radius of the inlet the relative diffuser length . The overall fenestron dimension is determined by the suction (low pressure) area on the intake surface. This dimension can thus be calculated according to the following formula .
The evaluation of the fenestron ability to generate thrust while consuming power is carried out according to the value of its energy conversion quality (power conversion factor) () [5]. This criterion was initially proposed by B. N. Yurev for determining the energy efficiency of the main helicopter rotor in hover mode.
The value of energy conversion quality of a fenestron in the International System of Units would be calculated according to the following formula   Here is the aerodynamic efficiency of the fenestron. shows by how many times the thrust of the «ducted fan» system is bigger than that of the open rotor. is the relative efficiency of the ducted fan.
The following equation is true for the hover mode of the helicopter at low altitude. Here is the specific thrust per unit of power (thrust efficiency). the specific thrust loading per unit of rotor disk area.
A fenestron with the following values of the design parameters was taken as an initial version for carrying out the numerical modeling: the fan hub radius , the number of blades 11, flat rectangular blades (with no twist) with NACA 23012 airfoil section, the gap between blade tips and duct walls 0,005R, the diffuser expansion angle 12°, the blade tip speed 220 m/sec.
The calculations resulted in the functional dependencies for the relative duct and fan thrust (which are taken relative to the total fenestron thrust ), the aerodynamic efficiency of the fenestron (), the relative efficiency of the ducted fan (), the energy conversion quality (power conversion factor) . All of the abovementioned characteristics were obtained as functions of the relative curvature radius of the inlet and relative diffuser length () . The calculations show that the rational values of  and  dimensions lie within the following ranges: .
Two fenestron design alternatives were considered: 1) a fan with the blades geometry, which provides equal speeds along the flow horizontal section at the diffuser exit; 2) a fan with flat rectangular blades and increased flow speed near duct walls. The calculations show the following: in the first case the fenestron fan has a higher relative efficiency () and lower aerodynamic efficiency due to the decreased induced power losses; in the second case the fenestron has a higher aerodynamic efficiency and lower value of due to an increased inlet (lip) thrust. At that the second alternative wins in terms of energy conversion quality and appears to be the preferable option. It was established that power losses on swirling the flow behind the fan amount to 7 8% of the power, which is consumed by the fenestron.
Thus it is recommended to install a flow-straightening device behind the fan.
The research demonstrates that the rational choice of the fenestron design parameters allows its power consumption, which is required for compensation of the torque reaction moment from the main helicopter rotor, to not exceed the power consumption of a conventional tail rotor by much.

Increasing of both flight velocity and load factor range by cargo detachment can be used for modern aircraft needed performances insurance. As the result both aerodynamic and inertial forces acting on cargo increase during cargo forced motion as well as cargo flight into turbulent zone at the carrier vicinity. All above mention should be considered especially applying to relative light cargo detachment.
Additional complexity of the detachment process arises because cargo in the carrier compartment has folded control surfaces. This causes to the so called effect of «deep well» which requires to performing detachment process by opening compartment doors. Moreover, existing of this «well» leads to an increasing of rigidity requirement both an aircraft and a device of forced detach structure. This additional requirement adversely affect to the overall weight of the complex.
Discussed above all mode of cargo accommodation into carrier compartment provides a set of new effects because cargo has folded control surfaces. At first, the internal cargo accommodation provides stealth of e quipped carrier but simultaneously increasing the height of the aircraft structure. At second, the internal cargo accommodation reduces aerodynamic drag forces of carrier but simultaneously increases the required path cargo forced movement during its detachment
Due to these regards, the current issue for safety of modern aircrafts is one of the topical problems. Authors suggest the method optimal criteria selecting in order to provide the safe cargo detachment from the internal compartment of the aircraft. This method ensures a guaranteed collisions absence by cargo detachment with compartment, aircraft internal elements, as well as doors and other cargo. The article consists also the results of various approaches to provide safety of cargo detachment, accommodate in carrier internal compartment.

Subject
The interplanetary trajectory of spacecrafts Article Optimal direct trajectories to the Jupiter with nuclear electric propulsion
Purpose
The analysis of optimal direct trajectories of spacecraft with nuclear electric propulsion during the flight to the Jupiter.
The analysis of the possible types of extremals.Using the transversality conditions to choose the optimal date of start and magnitude of hyperbolic excess velocity from the Earth
Design/methodology/approach
The trajectory optimization problem of spacecraft with nuclear electric propulsion to the Jupiter carried out by using the Pontryagins maximum principle.
Ensuring compliance with transversality conditions in solving the boundary value problem in order to choose the optimal date of start and magnitude of hyperbolic excess velocity from Earth.
Originality/value
The analysis of the possible types of extremals of the optimal direct trajectory to the Jupiter with nuclear electric propulsion
The transversality conditions in solution to the boundary value problem in order to choose the optimal date of start and magnitude of hyperbolic excess velocity from Earth.
Results
There are 8 types of extremals in rendezvous with Jupiter problem. The detail trajectory and mass characteristics of these extremals are presented. The analysis of choosing optimal date of start and magnitude of hyperbolic excess velocity from Earth is presented as well.
Practical implications
Space mission to the Jupiter with nuclear electric propulsion.
Conclusions
The comparative analysis between results of the using optimality conditions with the results of iterated date of start and magnitude of the hyperbolic excess velocity is presented. The coincidence of the results proves the correctness of used transversality conditions.
Analyzed space transportation system based on the heavy launch vehicle «Angara-5», the chemical upper stage KVTK and nuclear electric propulsion system with an input power of 100 kW can deliver a spacecraft with final mass 6246 kg in direct flight to Jupiter.

Purpose
The article is devoted to the technique of automated test of unmanned spacecraft flight plan. The technique is based on the use of spacecraft onboard control complex (SCU) as a test device. The execution of this test makes it possible to provide conclusions about the correctness of the flight plan and submit its progress during the flight to the SC. Design/methodology/approach
Inspection procedure of the created flight plan is based on the verification of a number of key parameters characterizing both the correctness of the flight plan transfer on the spacecraft and the correctness of its construction. It consists of three stages.
The transfer of flight plan on the spacecraft performed using formal structures called phrases. The first step is validation of the flight plan transfer provided as a set of phrases.
The second step is a test of acceptance of phrases in SCU. The integrity of the phrase, its structure and validness of data contained in it are checked. Next, the flight plan is converted to the type in which it was prior to its submission in the form of phrases.
The third test phase of the flight plan is in the process of passing through the SCU structure. It is performed with the technique called the «closure» condition. As a result of the SCU flight plan executing the data is generated and it is similar to content of the original plan that was sent to SCU. Comparing these data, we can determine the accuracy of the plan. If the plan was created correctly they will be equal. This check is the essence of the «closure» operation.
Findings
Feasibility check of spacecraft flight plan is an important step in the planning procedure which is impossible without SCU. In this case SCU must be presented in a form of a mockup or simulation model. The second option is more rational and has a substantial economic benefit. Considered structure of simulation model of SCU has shown that it can be applied to create models for different SCU SC. The proposed simulation model of the SCU was created and put into practice of SC control circuit. Its reliability and efficiency have been proven in normal and abnormal control schemes.
Research limitations/implications
The proposed method of checking the flight plan using a simulation model of the SCU can be used for flight control of unmanned and manned spacecrafts.
Originality/value
1. Checking method of flight plan consists of three phases. The most important part of it is the use of «closure» condition the essence of which is to compare the information specified in the flight plan and similar data obtained as a result of the passage of the plan through a simulation model of the SCU.
2. Using a simulation model of the SCU instead of the mockup can significantly reduce the check time of the flight plan to simplify the verification procedure itself and make it economically viable. 
3. Analysis of the structure of the SCU have showed similarity of performed function tasks for the various spacecraft. Because of this it is possible to create a basic framework for the development of such simulation models, as well as general requirements for them. 
4. The structure of SCU simulation model and function scheme of its blocks was created on the basis of the SCU «Meteor-3M» and «Coronas-Photon».
5. Presented data confirms the correctness of the SCU simulation model and proves the possibility of checking the flight plan correctness using this model. Keywords: spacecraft, flight planning, flight plan, control commands, simulation model, «closure» condition.

This paper considers a methodical apparatus for the rocket experiment designing; the peculiarities of problem statement for the rocket experiment designing and the algorithm of complex coordinated parametric optimization of technical means involved in a rocket experiment are discussed. Ratios for assessing RE efficiency at an early stage of development as well as comparative assessment of suggested RE facilities are presented.
Statement of research problem
Now in rocket experiments along with launching support means and systems (Launching Platforms, Operations and Checkout Complex, Ground Telemetry System, Command Position) a set of auxiliary systems is used to perform measurements and to transmit information.
The problem of rocket experiment designing in this case is associated with the assessment of major subsystem parameters, i.e. Research Rocket Complex parameters ground infrastructure parameters, parameters of means utilized to ensure the desired RE efficiency and minimum project realization costs.
Research technique
During the solution to the primary problem of RE designing the following scheme of partitioned iterative analysis of three main tasks is taken:
— problem of RRC parametric optimization at given parameters of ground infrastructure and utilized means;
— problem of parametric optimization of ground infrastructure at given parameters of RRC and utilized means;
— problem of parametric optimization of utilized means at given parameters of RRC and ground infrastructure.

The ratios for assessing RE efficiency are given. The efficiency of rocket experiment depends on the perfection of technical means, on the progress of tool- management and information-methodical support system. There are two groups of problems under consideration: one can be solved with the help of RRC, i.e. problems solved by direct rocket measurements, and problems solved by ground measurements of those effects in the atmosphere and space that are generated by devices installed on rockets.
Research results
Peculiarities of the algorithm of complex coordinated parametric optimization of technical means contributing to a rocket experiment are taken into account and the model of RE efficiency assessment was considered. During the development of such models the rocket experiment peculiarities are taken into account, experimental data as well as heuristic representations are applied.
Presented measures of RE efficiency are used to compare the efficiency of developed research rocket complex (RRC) MR-30 with its predecessor RRC MR- 12. With the help of innovations applied in its development the following characteristics are obtained: the delivery price of 1kg payload is one-half reduced; the body of obtained measurement information is increased twelve fold, moreover, the cost of information unit is decreased by more than twenty times.

Subject
The subject of this article examines the optimization of non-complanar multirevolutional low-thrust transfer.
Purpos
The purpose of this article is to demonstrate a number of new teaching techniques used in the formulation of the optimization problem, the definition of a mathematical model, as well as the approach for solution to the boundary problem using continuation method.
Methodology
The main methodological approach during the consideration of optimization problem is use of the Pontryagin maximum principle. Following the formalism of the maximum principle the optimization problem is reduced to the solution to boundary value problem. To solve the boundary value problem the continuation method is used. The quality criterion is the minimum engine time required to orbital transfer provided that the engine spacecraft is uncontrollable. In our mathematical model the equinoctial elements is used to describe the motion of the spacecraft. Right parts of the canonical system of optimal motion numerically determined in the process of integration. To determine the derivatives the complex step method is used. Also, the complex step method is used to determine the elements of the Jacobian matrix for the nonlinear system of the boundary value problem.
Results
The results of solution to the two types of problems have been analyzed. Two different orbital transfers from the original elliptical or circular orbit to GEO are submitted. The dependencies of determining the evolution of the trajectory and optimal control program for pitch and yaw channels are presented.
Practical implications
The results can be applied in the design of transfer vehicle with electric propulsion system. Also, these results can be applied in the guidance and control synthesis tasks for those systems.
Conclusions
The calculation of the right parts of the differential equations describing the optimal motion of the spacecraft in the integration process using the complex step method shown to be effective. This technique is possible to avoid the time-consuming procedures of analytical writing the right parts of the equations for the conjugate system.

Paper considers the Reusable Space Rocket System (RSRS) developed by the Khrunichev State Research and Production Space Center. The system includes Reusable Space Launch Vehicle (RSLV) and the ground complex facilities.
RSLV represents the multipurpose launch vehicle for the launch to the Earths orbits the automated spacecraft of various applications, manned cargo spacecrafts under the near Earth outer space exploration programs, for the Moon, Mars and other solar system planets exploration.
Basic element of RSLV is the Recovery Rocket Package (RRP). RRP aerodynamic configuration represents the design amalgamation of rocket (tanks, dry modules, liquid-propellant main rocket engine) and aircraft engineering (wing, vertical fin, tail unit, air- breathing propulsion system, landing gear) technologies. Thats why RRP is a very complex system, development and testing of which integrates not only rocket but the aircraft technologies and makes necessary the RRP and its components development test objectives and special features analysis. Paper shows that RRP and its components development test have the following special features:
— components development test is subsequently conducted with testing according to the components hierarchy level raising;
— during the RRP optimization the volume and depth of testing are limited, that is probabilistic because estimations of technical features based on the limited data amount are random;
— during the RRP development test the compatibility of systems and its components interference, and influence on operational capability of RRP can be evaluated only at the highest hierarchy levels;
— in the course of RRP development test its components technical features testing can be nondestructive or destructive. It leads to the variation of testing costs from tens of thousands rubles for the low level components nondestructive testing to millions and hundreds of millions rubles for destructive testing of finished products;
— products manufactured in present time by aerospace and aircraft industry enterprises have sufficient continuity of structural schemes and used technologies, mostly they directly borrowing some elements of equipment, systems and etc. from earlier realized projects. This statement is the basic point for qualification method which allows to avoid excessive
— during the development test the specific decisions are taken at each hierarchy level, it could be right or wrong because of the limited depth of testing and various level of technical parameters measurement instrumentation errors, and the consequences of wrong decisions are different for the specific levels.
All the foregoing particular issues arising during the RRP parts specifications control at the stage of development test should be considered in their testing plan at every level. At the same time, present methodical support for the development test does not provide this opportunity and it leads to the necessity for development of the new mathematical models, algorithms and RRP parts engineering data hierarchical control of optimization methods.

Configuration of magnetic field in the discharge chamber of ion thruster is traditionally considered in the axisymmetric approximation. This approach is reasonable for the thrusters with permanent magnet rings (ring cusp thruster). However, as it is shown hereinafter, the Kaufman thrusters using magnetic systems with electromagnets placed around the discharge chamber periphery have significant azimuthal nonuniformity of magnetic field. This nonuniformity can lead to the reduction of magnetic field capability to confine electrons from direct loss to the anode. There are two reasons of such reduction. First, magnetic field has too low strength in the regions between magnetic coils to confine electrons efficiently. Second, new mechanism of electron drift appears in the nonuniform magnetic field.
In this paper, two magnetic system configurations of the ion thruster IT-300 are investigated. The first one has 12 coils with the current I = 3 A, the second one has 6 coils with the current I = 6 A. According to the magnetic field measurements, such configurations have the same average magnetic field strength (~3,7 mT) and different degrees of azimuthal nonuniformity.
Laboratory tests were conducted for two investigated configurations for the purpose of determination of radial and azimuthal distribution of the beam current density. This distribution is characterized by the nonuniformity coefficient being a very important parameter in terms of the thruster lifetime. Measurements were carried out using five Faraday probes. The probes were placed on a mounting bracket so that the distances between the accelerating grid and collector of each of the probes were the same. Collector was made of molybdenum and had the collecting surface diameter of 10 mm.
Tests to determine relationship between the discharge loss and mass utilization efficiency were conducted for the estimation of discharge chamber operation efficiency. Tests were carried out with the accelerating voltage of 2500 V and beam current of 200 1400 mA.

In the article, the methodology and algorithm of the process of deformation of the shell under the action of dynamic pressure in general case of a variable over the shell surface are considered. This algorithm can be applied for the shells of rotation of an arbitrary shape. Such problems appear during the shaping of the shell- type structures from the flat plates.
As a research method, the finite elements method is applied. The system of differential equations is solving by Runge-Kutta method with the automatic selection of the loading step to the specified accuracy.
The developed methods and algorithms allow analyzing large displacements of thin shells and plates under the dynamic loads. The solutions of the problems of large displacements and deformations are fall into several stages. Stiffness and vector of generalized external forces are recalculated when the deflections are reached the order of thickness of a plate. The recalculation is based on the achieved level of stress-strain state of the shell and changes in the geometry of the shell.
The algorithms were used to calculate the shaping of a conical shell from a circular plate under dynamic loading. Accordance of the theoretical and experimental results observed.
The main result of the study shows that it is possible to control the loss of stability of thin plates and shells at high plastic strains under the dynamic loading.

One of the ways of representing characteristics of fatigue resistance of materials and structural elements is the distribution function of the limit of endurance. It order to calculate structural elements, to conduct test planning and to perform statistical analysis of their results it is recommended to use the normal law of distribution , of the limit of endurance or the normal variable distribution (logarithmic normal where m — the number of levels of voltage amplitude distribution) [2, 3]. For steels Weibull-Gnedenko distribution of the limit of endurance is widely used. Statistical estimation of the parameters of the distribution function of the limit of endurance is performed by «up- and-down» and «probit» methods based on the maximum likelihood method (MLM), as well as by other methods.
According to the «up-and-down» method [19], the first sample of a series of n objects is tested at the expected level of endurance. If the first sample is not destroyed to the basic number of cycles , then the second sample is tested at a higher voltage, and if it is destroyed, then the second sample is tested at a lower voltage. Voltage level for testing the third sample is chosen depending on the test results of the second sample.
MLM-estimations of distribution parameters are defined by solving the system of (according to the number of parameters) equations:
where m — the number of levels of voltage amplitude cycles during the tests;
the number of destroyed objects at i-th level;
the number of undestroyed objects at i-th level;
the number of tests at i-th level;
the total number of tested objects;
  probability of destroying at i-th level of amplitude;
сontinuous differentiable function of the level of endurance distribution with parameters g, which are to be estimated according to the theoretical law of the level of endurance distribution.
When tested by the «up-and-down» method, are random variables.
Derivatives define the specific form of the system of equations (1).
According to the «probit» method [13], a series of n samples is divided into 4 or 5 groups. Samples fr om each group are tested to the basic value of the number of cycles at the appropriate voltage level. As a result of the tests there appear destroyed and undestroyed objects at each level. Estimations of parameters of the endurance level distribution are defined by solving the system of equations (2). Covariance matrix is defined by the equation (8). As opposed to the «up-and-down» method, the number of levels of voltage amplitudes m, as well as the number of tested samples at each level n_i, are defined in advance based of the test plan.
The system of nonlinear equations of maximum likelihood (1) for estimation of parameters of normal, logarithmic normal and Weibull-Gnedenko distributions has, as a rule, several local minimums. That is why any standard method of numerical solution of the system of nonlinear equations will not be able to provide a single- valued solution, as the solution will depend significantly on the initial approximation. In these circumstances the primary task is, first, to establish exact approximations in order to justify initial approximations, and, second, to select the most acceptable numerical method of solving the system of nonlinear equations.
In this paper we recommend to use the method of deformable polyhedron (by Nalder-Mead) in order to solve this system, which, in practice, makes it possible to find the roots of the system of nonlinear equations rather fast and with the required accuracy. The authors have developed the program of implementation of this method (when writing the module, the authors used the data of the cross-platform numerical analysis library ALGLIB according to license terms GPL 2+).
Conclusions
1. It has been established that the system of maximum likelihood for estimation of parameters of normal, logarithmic normal and Weibull-Gnedenko distributions of the level of endurance at fatigue tests with the use of «up-and-down» and «probit» methods has, as a rule, several local minimums which makes the numerical solution difficult, as the solution will significantly depend on the initial approximation.
2. In these circumstances there has been developed an algorithm and computer programs for calculating initial approximation and subsequent numerical solution
3. The algorithm developed by the authors allows fast and accurate calculation of estimation values and lower confidence bounds of the limit of endurance at fatigue tests.
The approach, described in work is especially actual in problems of calculation of a limit of endurance of the various expensive designs meeting in aviation systems and systems of rocket and space equipment, wh ere a large number of mechanical tests is economically inexpedient.

This paper discusses the use of situational modeling for the control process, target distribution of a group of unmanned aerial vehicles (UAV) during the interception of air targets in the attacking mode. General block diagram of the control system of situational type is presented. The method of situational modeling is based on three principles. The first one supposes the possibility of situation classification. The second principle is based on the idea of using special formal means (languages of knowledge presentation) while describing the object under control, the processes of its functioning, and the environment. Such means secure completeness of description for the objects under control and equivalence to the real process. The third principle supposes building the model of object control or the model of entering solutions.
Examples of resultant rules used for the description of target distribution algorithm are presented also. They allow adequate description for tactical situations using numerical and logical values, and building a decision tree of the target distribution algorithm. The prediction unit comprising extrapolator is the most important part of the target distribution algorithm. This unit provides forecasts for the behavior of the enemy and of the friendly UAV, simulation for all possible options of target distribution, as well as obtaining of the simulation result for the certain tactical situation and selection of the best option based on the given criterion of efficiency of this algorithm. The extrapolator includes the models of UAV and targets, and comprises a unit to control such models. Mathematical model for the UAV description is presented, which includes dynamic and kinematic equations of the mass center motion and kinematic equations of motion relative the UAV mass center in case of the gliding angle and side acceleration . The presented mathematical model allows adequate description for the controlled objects (UAV and targets) and the control processes proceeding in the objects of control. The criterion of efficiency evaluation is formulated for the suggested target distribution model. It characterizes the number of targets shot down during the minimum attack time. An example of efficiency evaluation for the target distribution algorithm is presented also.

The purpose of this research was to develop a computational-experimental technique for remote (non- contact) diagnostics of structural defects in elastic materials. The performance of equipment is based on nonlinear interaction of acoustic beam of finite amplitude and analyzed material. Elastic properties of flaws differ from ones of base medium. Gradient of elastics properties on the border of flaws results in the appearance of non-classic nonlinearity in the domain. Such nonlinearity significantly exceeds the physical nonlinearity of a medium under diagnostics. Determination of spatial distribution of structural acoustic nonlinearity in the specimen allows to spot the defects.
The most promising direction in further development of non-destructive diagnostics (research methods) for the elastic composite materials is to use the procedure of inverse problems for estimation. Such problems are of great practical importance in the study of properties of composite materials used as non-destructive elastic surface coating in objects of space technology, power engineering etc. The developed method could be applied for determination of materials properties; the availability of corresponded experimental facilities allows us to provide uniqueness of the solution.
The iterative regularization method is outlined as one of the most efficient and universal ones for solving the ill-posed inverse problems, which arise in the course of the diagnostics. Therefore an extreme method of solving of ill-posed problems of wave propagation is developed. This method is based on the iterative regularization principle. The stability of the solution is achieved by timing the number of iterations in the gradient methods of the first order with the error of initial data. In the beginning rigorous mathematical results are stated in brief substantiating of this approach with respect to the most widely spread gradient algorithms, i.e., steepest descent and conjugate gradients, namely, information is supplied on regularization of the mentioned methods by the number of iterations and legitimacy of using or residual criterion to choose the number of the last iteration or restart the iterative process. To solve the inverse problems there have been developed iterative computational algorithms, implementing a minimization through gradient methods of residual functional, characterizing the square deviation of pressure values computed and measured by means of the assumed mathematical model. Here, the unknown functions are approximated by the cubic B-splines and the inverse problem is thus reduced to a definition of the unknown parameter vector. The gradient of residual functional is computed using the solution of boundary-value problem for the adjoin variable. In the approach under consideration all experimental information is used to estimate material characteristics as pressure functions on the all pressure interval of interest.

Changing the innovative production creating processes on aviation enterprises due to innovation model improving. Widespread is the new innovation model, initiators and sources of which are the users of innovative developments — industrial complexes, enterprises. This requires the adequate management tools development, which have sufficient organizational, financial, institutional and spatial flexibility.
The work object is to distinguish enterprise management features and life cycle processes of innovations creating, associated with the ability to get information resources fr om the environment and transfer their knowledge / products / technologies to consumers.
The life cycle innovation model, proposed in this work is based on the operations research theory and the complex aerospace systems theory.
The research results are:
- A management process model for life-cycle stages of innovative products creating, including the directions and works sequence upon the proposed stages;
- Marketing appraisal parameters and selecting the innovative idea application system;
- The sequence of organizational and management directions of innovative ideas search and selection, formation of ideas appraisal criterias system and requirements to the developers;
- Organizational and technological processes directions during product samples manufacture or the product manufacture pilot lot;
- The key competences selection at the testing and commercialization stage, based on risks and experimentation continuing benefits appraisal;
- Organizational and management processes of innovative products / technologies introduction and dissemination on markets.
The scope of the research results application are management systems of high-tech industrial complexes, which initiate innovations, as well as enterprises, operating in the initial phase of the innovative products development and manufacture and / or when entering a new, incipient market.
Summary 1. Use of the innovation model, where a developer is the industrial complex and enterprises included in it accompanied by transformation of management and organizational processes of units production, information environment, personnel management and selection processes. 2. The innovative production programs development managing at first requires the processes formation by stages of units production life cycle, wh ere are various aspects, as marketing, research, production, technological, financial, economic, etc. 
3. The proposed development is essential to improve the innovative enterprises intercompany activity in the aviation industry in horizontal linkages activation by stages of life cycle, allowing to improve users practice and further commercialize the idea.

There exists a number of difficulties by placing guided missiles (GM) into fuselage of the carrier related to the limited space and tight GM placement. These circumstances may cause increased demands on the GM bailout system which should provide GM bailout process without touching GM body and control surfaces both carrier compartment elements and other missiles. An important feature of GM separation chamber is GM length of flight in a limited space due to the substantial depth of the compartment, which leads to a rather large time of the not stabilized GM flight.
Examination of the mentioned above collisions eliminating methods is performed on the basis of the two-stage type ejection device (ED). The first stage of ED prints a second stage with a suspended GM from the carrier compartment to the border of the fuselage and locked it in the down position. GM bailout is performed by the second step of the ED, which can be considered as elastically suspended to ED first stage,fixed in the lowest position.
Description of the system dynamics is based on the second kind Lagrange equations utilization. Using comprehensive mathematical model of the relative GM and carrier motion has been performed simulation of both launching and initial portion of the GM flight applying to the carrier wide range flight conditions Simulations showed that the pre-installation (before the bailout) to dive rudders GM with the values formed safe relative trajectory of AGM in all modes .

Purpose
The research was aimed at developing a method of assessing the effectiveness of usage of multi-purpose aircraft systems and individual units (MPSU) on the advanced military aircraft during early stages of their design.
Design/methodology/approach
The paper adduces the analysis of the possible ways of using MPSU on the advanced types of military aircraft.
The further increase of the aircraft efficiency depends on attaining several conflicting objectives in the field of aerodynamics, layout design, combat survivability, reliability and operation of aircraft. This in turn induces in a wide search for new design solutions. The separate further improvement in each of these individual areas is often ineffective due to the high degree of systems integration and interconnection. One of the most promising ways of solving these conflicting problems together as a complex consists in the application of design methods, which ensure the multi-purpose usage of individual units and systems of the airplane.
The groups of aircraft parameters, which are significantly influenced by the usage of the MPSU, were identified and structured. Partial and integral criteria of effectiveness of aircraft with MPSU were allocated.
The research included identification of the areas, in which the usage of the multi-purpose systems and units is the most relevant for improving the aircraft performance. Such areas include aircraft structure, systems and equipment. The control systems were identified as the most promising area for MPSU usage. In particular, it is advantageous to install the rotary actuators, which would act both as actuators and load- carrying structural elements of the flight control device, inside the aircraft control surfaces. Such solution creates additional space for extra fuel tanks in the fixed part of the wing, reduces the weight and moments of inertia of the control system and improves the aircraft maintenance characteristics.
The research helped to determine the main parameters in various areas of aircraft design, which are influenced by usage of MPSU most of all.
The results of the work allowed to compile a flowchart diagram, which presents the developed methodology for MPSU effectiveness assessment. This diagram shows the application of both consecutive and parallel methods of determination of the integral criterion of economic effectiveness, which are based on simultaneous definition of the required parameters in various design fields.
Thus, by using this approach it is possible to assess the effectiveness of the proposed versions of multi- purpose systems and units within the structure of the advanced aircraft already at the early design stages. Findings
The research resulted in the development of a new approach to the problem of assessment of the effectiveness of usage of multi-purpose individual units and systems on the advanced military-aircraft. Originality/value
Application of the developed method in aircraft design will allow to reduce the required costs and time as well as increase the quality of the results of design activities.

The article suggests use of utility function theory for assessment and decision making in the process of best conceptual design selection of the aircraft of the military- industrial complex (MIC). The utility function allows combining of manufactures and consumers approaches within the process of new technics development and in the same time solves task of the best alternative selection among considered projects. For manufacture the utility is a mean to assess the degree of the product being attractive to consumers and for potential customer utility is a tool to sel ect best alternative among considered.
The usefulness of an alternative for the consumer may be achieved with some probability (stochastic case). If selection of the alternative brings certain utility with certainty it is considered deterministic case. The first case in the literature is referred as theory of utility function, the second as theory of value function. In our analysis we concentrated on deterministic case.
Presented theory of cardinal value function gives the framework to form the set of criteria describing alternatives of conceptual design. The cardinal feature of the value function makes it possible to determine not !«only order of the alternatives but also how far one alternatives stands from one another in terms of utility. This property allows use of the results of the decision making process in investments decisions. One can form the order of the alternatives having their utilities and dividing them by investment needed for the project. Such order shows amount of utility fr om the project per capital invested.
The axioms are given, on which existence of cardinal value function is based. Key definitions and theorem to support forming additive multi-attribute cardinal value function have been outlined.
The task of designing of the technical system consists of two parts — generation of the alternatives and selection best alternative. In the example it is considered that the first part of the designing process is solved and it is considered multi-attribute decision making problem of selection of the best alternative. The example demonstrates use of the cardinal value function theory for selection of the conceptual project of the surface to air missile (SAM) among alternatives that have different competitive advantages. The major design objective is to select the design of the missile that destroys given types of targets better than the other projects. The value function has been built based on three criteria that characterize project of SAM: the targeted efficacy, time to capture, and initial mass. The method does not have limitation on number of criteria to include in the value function. In order to support possibility of building additive value function, it is briefly discussed difference independence of the mentioned three criteria. The process of obtaining single attribute value functions has been omitted due to technical nature of the process putting together experts opinions on attitude to desirable and non-desirable particular values of selected criteria.
The suggested method can be used in the automatic decision making systems as the assessments on conceptual design can be obtained «on the fly» as soon as values of the selected criteria become available.

The paper is focused on the research of the influence of the diffuser length on the aerodynamic characteristics of the «shrouded rotor» system. «Shrouded rotor» is used as lifting and propulsive system on many types of aircraft. Single-rotor helicopter fenestron with a limited diffuser length, which is built into the tail boom fin, is an individual version of such «shrouded rotor» system. The early works of the author adduce the methods of calculation of aerodynamic characteristics of the shrouded rotor with a diffuser, which is long enough to provide the full expansion of the exit air flow stream. However, the fenestron exit flow cannot fully expand due to the limited length of its diffusor. Thus there is a contraction in the exit flow stream of the fenestron, which affects its aerodynamic characteristics.
The adduced calculation method considers the «shrouded rotor» system as a complex of elements, which create hydraulic drag in the passing air flow. The complex consists of an inlet and a short diffuser. These elements compose the duct where the rotor and the internal devices, which support the rotor and create local pressure losses within the complex, are installed.
The rotor itself is modeled according to the scheme of the rotor disk theory with taking into account the additional induced flow, which is created by the shroud. It was established that the relative curvature radius of the inlet and relative diffuser length (where R is the shroud radius) influence significantly the improvement of the fenestron aerodynamic characteristics, which is attained due to the aerodynamic force performance of the inlet at the duct entrance. Shroud lift effectiveness is determined relative to the total system thrust .
Calculations show that to attain the best shroud lift effectiveness its design parameters should be sel ected fr om within the following ranges: ; .