Научная электронная библиотека ULRICH'S Periodicals Directory

Сибирский государственный университет
науки и технологий имени академика М.Ф. Решетнева

Сибирский аэрокосмический журнал
ISSN 2712-8970 (Print) ISSN 2782-5760 (on-line)

Сведения об образовательной организации

UDK 621.454.2
THE METHOD OF CALCULATION OF THERMAL STATE OF CHAMBER FOR STEADY STATE IMPULSE OPERATION OF LIQUID ROCKET ENGINE OF SMALL THRUST
A. G. Vorobyev*, S. S. Vorobyeva
Moscow Аviation Institute (National Research University) 4, Volokolamskoe sh., A-80, Moscow, 125993, Russian Federation
In this article, there is a method of calculation of thermal state of chamber of liquid rocket engine of small thrust. There is a physical picture of the processes occurring in the small thrust rocket engine during start and stop moments of motor. The difficulties in modeling the processes during star and stop of small thrust jet engine are presented. The cause is conjugated intra-chamber physically processes occurring at varying parameters during the mixing elements in the off-design conditions. In the article, a mathematical model to predict the thermal state of the combustion chamber for the different modes of engine operation is presented. To simulate the process of starting and stopping the engine applies a quasi-steady approach that replaces the transient process by set of steady-state processes with variable time operating parameters. The initial data for mathematical model is parameters of small thrust rocket engine, developed in Moscow Aviation Institute, working on the components of the nitrogen tetroxide and unsymmetrical dimethylhydrazine, thrust 200 N. The engine is equipped with a separated mixing head, which contains two-component jets, and film cooling by fuel that protects the combustion chamber wall. The results of mathematical simulation of the small thrust rocket engine in the steady pulse modes are given. The pulse modes at 0.05 sec and pause at 0.05 sec, pulse mode at 1 sec and pause at 1 sec and mode, consisting of sets of pulses with different duration are analyzed. The wall temperature of combustion chamber in characteristic sections, depending on the wall temperature by time for internal and external surfaces of the combustion chamber is calculated. The results of calculation for impulse mode approve a large temperature gradient on the internal wall surface of the chamber between pulses are shown.
Keywords: liquid rocket engine of small thrust, rocket thruster, thermal state, mathematical modeling, steady pulse mode, nonstationary thermal mode, combustion chamber, film cooling.
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Vorobyev Alexey Gennadievich – Cand. Sc., docent, Department of Rocket engines, faculty of Aircraft engine,

Moscow Aviation Institute. Е-mail: formula1_av@mail.ru. 

Vorobyeva Svetlana Sergeevna – senior teacher, Department of Rocket engines, faculty of Aircraft engine,

Moscow Aviation Institute. Е-mail: kinder-svetiks@yandex.ru.