Abstract
In the combustor for rocket engines, liquid film cooling is a widely adopted technique for restricting the structural components to the acceptable bounds for the successful operation of the propulsion systems. Multiple cooling techniques for thrust chamber walls are favored along with the coolant film in propulsion systems operating at higher chamber pressures by incorporating an ablative cooling mechanism for the nozzle. The adoption of combination will result in challenges in simulation studies due to the presence of multiphase phenomenon in the system. The current article presents the effects of these two cooling methods on a thrust chamber wall studied through compressible multiphase formulations. A majority of the previous studies have used incompressible flow equations to model coolant film behavior and associated heat transfer. The present study utilizes an approach utilizing compressible multiphase flow to simulate the behavior of the compressible hot gas flow and the incompressible coolant liquid film accounting for phase change effects, vapor diffusion into hot gas, radiation effects on coolant surface, liquid entrainment, and blowing effects due to vapor formation at interface. Film-cooling effects on ablative nozzle accounting for pyrolysis phenomenon due to material decomposition governed by Arrhenius expression are also emphasized. The outcome of the numerical model showed good agreement with available test data in literature, and results from the tests done in-house. The model described in the present study was able to support the actual evidence of liquid film cooling being able to preserve the walls of the thrust chamber from severe internal thermal environment and prevent ablation on surface of nozzle.
