Thermal Analysis of Cryogenic Rocket Thrust Chamber with Regenerative and Film Cooling

Abstract

The thrust chamber is a critical component of the rocket engine. It converts the chemical energy stored in the Propellant into Kinetic energy required to generate the necessary thrust. The thrust chamber consists of an Injector, a Combustion chamber and a Nozzle. The propellant (fuel and oxidizer) is injected and ignited in the combustion chamber, generating high temperature and high pressure gases. This energy of combustion heats up the reaction product gases to a very high temperature, typically to 4500 K. This in turn raises the pressure inside the combustion chamber to very high values. These product gases enhance the heat flux of the chamber wall, which produce thermal stresses and decreases the chamber life. In order to improve the life of the chamber, cooling techniques become a major design consideration. In this study, the influence of mass flow rate of film coolant gas has been analyzed while keeping the mainstream inlet condition unchanged. CFD analyses have been used to understand the behavior of various factors such as curvature, aspect ratio, mass flow rate and thermal behavior in a Regenerative and Film Cooled Cryogenic Upper Stage Engine of the GSLV. Also, its 1-D mathematical calculation is done in MATLAB. The propellants used are cryogenic LOX and LH2. The chamber inner wall consists of copper due to its high thermal conductivity and the outer wall is designed by Stainless Steel because of its high strength.