Numerical Study of Reacting Gas Flow Heat Transfer in Liquid Rocket Engine Thrust Chamber

In this article, the reacting gas flow heat transfer is numerically studied in the thrust chamber of a liquid rocket engine. The gas flow in the thrust chamber is considered as a quasi one-dimentional mathematic model in which the effects of friction, the convection heat transferred to the wall, gas radiation, and non-equalibruim chemical reactions were taken into account. The gas governing equations are solved with a fully implicit finite volume method. The Brabbs air-hydrogen chemical model, which includes 9 species and 18 finite rate elemntary reactions, is used for modeling the combustion processes. The coolant flow in the cooling channels and the heat conduction through the wall are modeled as a quasi one-dimentional mathematical model. Numerical results show that the method used for solving the governing equations can appropriately simulate the gas flow field, the wall temperature, the heat flux through the wall and can calculate the increase in the coolant temperature in the cooling channels, as well. Besed on these results, the wall temperature and the heat flux through the wall on the hot gas side reaches a maximum aboat the throat. The maximum heat flux by gases radiation is about 30% of total heat flux to the wall of the combustion chamber. Comparing the results obtained for this code that is considered as the effect of chemical reactions in the chamber with the numerical calculations, it can be seen that the differencess in the maximum heat flux to the wall, the maximum wall temperature, the tempreature increase, and pressure drop of the coolant in the cooling channels are 30%, 7%, 29%, and 3%, respectively.