Evaluation of the thermal state of the combustion chamber rocket engine of small thrust using environmentally friendly propellants

Аuthors

Kovateva Y. S.^*, Bogacheova D. Y.^**

Moscow Aviation Institute (National Research University), 4, Volokolamskoe shosse, Moscow, А-80, GSP-3, 125993, Russia

*e-mail: kovateva2005@rambler.ru
**e-mail: bogachulya@mail.ru

Abstract

Study the possibility of using existing engineering design procedures (EDP)[1,2,3] assessment of the thermal state of the combustion chamber of liquid rocket engine (LRE) in relation to the rocket engine of small thrust.
The comparison of the calculation of results thruster with gaseous oxygen and gaseous methane, which the simulation process in ANSYS CFX, and the results of calculations by well-known methods are made. We consider two ways to determine the mixing flow with the wall layer. The first method is based on the calculation of the turbulent mixing flow with the wall layer by engineering methods of calculation for LRE of large thrust. In this method, the intensity of mixing depends on the chosen value of the constant coefficient K, varying from 0.05 * 10-2 ... 0.2 * 10-2. The second method is based on the use of ANSYS CFX for workflow of mixing ingredients without combustion by SST turbulence model with the value of the turbulence intensity equal to 5%.
We consider two ways to determine the level of heat flow and temperature of the wall. The first method is based on methods that are applicable for the calculation of heat flows for LRE of large thrust. The second method is based on application of the distribution ratio of the components along the wall with the data obtained in ANSYS CFX.
When comparing the engineering method of calculating the results of mixing and mixing flows ANSYS CFX we received a significant difference in areas of complete mixing. Thus, it calls into question the use of methods of calculation of mixing flow with the wall layer applicable to LRE of large thrust for the calculation of processes in the chamber of rocket engine of small thrust. Further study and correction values of K are required.
Calculation results for the heat flow and temperature level the outer surface of the combustion chamber two ways shows qualitatively similar values ​​in satisfactory agreement with the experimental data. This proves the applicability in the calculations of any technique.

Keywords:

rocket engine of small thrust, gaseous oxygen gaseous methane, engineering method for calculating, numerical simulation, fuel ratio, wall layer, film cooling, temperature of combustion products

References

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4. Kovateva Y.S, Vorobiev A.G., Borovik I.N., Khokhlov A.N., Kazennov I.S. Zhidkostnyj raketnyj dvigatel' maloj tjagi na toplive gazoobraznyj kislorod i gazoobraznyj metan (Liquid rocket engine of small thruster with fuel gaseous oxygen and gaseousmethane), Vestnik Moskovskogo aviatsionnogo instituta, 2011, vol.18, no. 3, pp. 45-54

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