Investigating methods of determining the thermal loading of the aircraft during launch sequence


Аuthors

Denisov M. A.1, 2*, Medved V. S.1**, Ludilshchikova S. I.1

1. Tactical Missile Corporation, 7, Lenin str., Korolev, Moscow region, 105005, Russia
2. Bauman Moscow State Technical University, MSTU, 5, bldg. 1, 2-nd Baumanskaya str., Moscow, 105005, Russia

*e-mail: mikdenisov@gmail.com
**e-mail: mvbarny11@gmail.com,

Abstract

The development of advanced aircrafts is one of the priority tasks of the aerospace industry. At the stage of experimental design work (EDW) for the creation of aircraft, a wide range of science-intensive and technical problems of varying complexity are solved, including the calculation of the temperature regime of its structural elements. The study of methods for solving problems related to thermal loading opens new opportunities for optimal design of aircraft, rocket and space equipment.
High heat flows acting on the product during operation, together with high aerodynamic loads lead to the need to strengthen the structure of aircraft, using active and passive heat protection, increasing mass characteristics and reducing the characteristics of aircraft. For this reason, the correct assessment of the thermal state of the aircraft at the early stages of development work significantly affects the design decisions taken in the development of products, the cost and duration of design.
The following computational and experimental studies have been carried out to achieve the goal:
1. The problem defining the heat flow to the aircraft during flight using the method of criterion relations and direct calculation of the coupled problem have been solved.
2. Calculations of the thermal regime of the aircraft components were carried out;
3. Comparison of the investigated calculation methods has been carried out;
4. Validation of the results of calculation studies of the thermal mode with the results of experimental work on the test site using a full-size sample of the product was carried out.
Using the criterion and coupled approaches, calculations of the temperature regime at the stage of the aircraft launch with the description of the temperature field during further operation of the product were performed.
The results of cross-verification analysis confirmed qualitative and quantitative convergence of maximum temperatures in the control points of the aircraft design. In the temperature range from plus 200 °C to plus 400 °C the discrepancy does not exceed 20 °C, in the range up to plus 80 °C - 4 °C.
By results of researches qualitative and quantitative convergence of methods of definition of boundary conditions with use of criterion equations and direct modeling of the flow over the first compartment of the aircraft is confirmed. Absolute deviation of temperature determination obtained by comparing temperature probe readings and results of direct calculation does not exceed 3 °С in the temperature range up to plus 60 °С. The temperatures obtained by the results of calculation using the criterion relations are overestimated - the absolute deviation between the calculation results and the experimental data does not exceed 7 °С with the experimental data.

Keywords:

validation, experiment, aircraft structure, heat exchange, aircraft, mathematical methods

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