Abstract

Abstract:- The present analysis deals with the estimation of ablation of thermal protection system provided at the inner surface of rocket nozzle using finite element method. The temperature due to hot gases varies from point to point in convergent, throat and divergent portion of the nozzle. The previous studies on ablation of rocket nozzle liner, assumed uniform variation of heat across certain portion along the axial direction. In present study the solution is refined for actual distribution of thermal load at the inner surface of the liner in nozzle. It is observed that, in the present case, the predicated ablation is less when compared to the previous studies and hence recommended for optimum design of liner