Liquid film cooling is commonly employed in the combustor of a liquid rocket engine to restrict the wall temperature within allowable limits. In case of few rocket systems, additional cooling is carried out by including an ablative nozzle throat section, apart from liquid film cooling. In the present study, applications of both liquid film cooling and ablative cooling methods for a rocket thrust chamber are presented. Most of the earlier studies have considered incompressible flow approach to simulate the film cooling heat transfer. Here, a numerical model based on compressible multiphase flow approach is used for heat transfer between the hot combustion gas and the liquid film injected along the wall. Phase change of liquid and diffusion of vapor into the gas flow are incorporated, to investigate the effectiveness of cooling along the solid wall. Variations of liquid film, temperature and mass fraction and their effects on wall-cooling, are studied. Ablative cooling is considered for the nozzle throat region and the heat transfer is studied considering the material degradation, including heat conduction to the wall. The results obtained from the numerical model compare favourably with experimental results available in literature, as well as data obtained from in-house engine tests. Liquid film cooling is observed to protect the thrust chamber from direct exposure to hot combustion gas and also eliminate ablation of the nozzle throat insert.