Numerical Analysis of Combustion and Film Cooling in LOx-Methane Rocket Engine

Nowadays methane is considered as an alternative fuel for rocket engine application. It is abundant in the outer solar system that can be harvested from mars, Titan, Jupiter, many other planets and moon. The properties of cryogenic methane like higher density, higher vaporization temperature, less challenge in storage requirement compared to cryogenic hydrogen and its higher specific impulse, superior cooling capability and higher coking limit compared to kerosene created a renewed interest on space scientist in choosing methane as a propellant. Combustion and film cooling involving supercritical and transcritical regimes in a high pressure rocket engine combustion chamber is a challenging problem for the rocket engine designers. At supercritical and transcritical conditions, the thermophysical and transport properties exhibit tremendous variations which makes modeling combustion and cooling a demanding task. The present study is based on the oxygen methane thrust chamber model B at DLR P6.1 test bench. The finite volume based software ANSYS Fluent 15.0 is used for the numerical studies. The combustion is simulated using single step eddy dissipation approach. The real gas effects are incorporated using Soave Redlich Kwong cubic equation of state. The numerical scheme is validated by comparing the results obtained with the experimental results reported in literature. Investigations have been carried out to study the effect of film cooling on chamber wall temperature and wall heat flux for different oxidizer inlet temperature. The result revel that the hot gas temperature near the wall is higher in the transcritical combustion compared to supercritical combustion.