Experimental Investigation of Adiabatic Film Cooling Effectiveness and Heat Transfer Coefficients Over a Gas Turbine Blade Leading Edge Configuration

Increasing the rotor inlet temperature is one of the key technologies in raising gas turbine engine performance, for which the turbine blades need to be cooled. Film cooling is one of the efficient cooling techniques to cool the hot section components of a gas turbine engines. In film cooling, a gas which is cooler than the main stream is passed onto the external surface via small slots or rows of holes within the surface. In the present study, the experimental investigation was conducted for an adiabatic film effectiveness and heat transfer coefficients over a gas turbine blade leading edge model at a subsonic cascade tunnel facility of CSIR-National Aerospace Laboratories, Bangalore. This study was aimed to investigate the effect of blowing ratio on the adiabatic film cooling effectiveness and heat transfer coefficients experimentally for the 20 degrees hole inclination angles gas turbine blade leading edge model. The blade leading edge model was fabricated using the rapid proto typing method using a very low thermal conductivity nylon based alloy material. This study aims at bringing the optimized blowing ratio values for the considered hole diameter of leading edge configuration. The comparative results showed that the blowing ratio beyond 2.0 does not have any improvement in the adiabatic film cooling effectiveness.