Abstract:
An experimental investigation was carried out to determine the extent of reduction in wall temperature in a model thrust chamber of liquid rocket engines for
various configurations of film cooling. The experimental test rig comprised a 100 kW hot air supply system, test section and required instrumentation. Hot air simulates the hot combustion products and nitrogen gas is used as the film coolant. Coolant gas injectors of three different configurations, viz., straight and compound angles of 300-100 and 45°-10°, were used for the experiments. Experiments were carried out for different core gas flow temperature over a range of blowing ratios. The study brought out the relative merit of each of the injector configurations with respect to film cooling effectiveness, film cooled length and its uniformity. The main-stream hot gas used at two different temperatures of 343 K and 397 K with flow rates of 0.13 and 0.094 Nm3/s for both temperature and the coolant gas applied at 308 K with flow rates of 0.0028 and 0.0021 Nm3/s. Initially, the observations of surface temperature of test section were made with the help of thermocouples without applying the coolant and then with coolant for different coolant injections. The data that are available from the experiments is used to evaluate the film cooled length, film cooling effectiveness, film stability and the effects of blowing ratio. It is also used to compare the above evaluations for all configurations and hence to conclude the effect of the configurations on film cooling. It was noticed that the film cooling effectiveness was higher in the case of injector configuration with compound angle injection of 300-100 as compared to that of straight and the compound angle of 450-10° injection configurations. The film cooled length was also found to increase with blowing ratios. Higher core gas temperature lead to a reduction in the effective film cooled length.
