Heat Transfer Near the Entrance to a Film Cooling Hole in a Gas Turbine Blade

Abstract

Film cooling is a method used to prevent jet engine turbine blade failure due to overheating. It consists of bleeding relatively cool air from the engine's compressor stage and discharging it through small holes in the turbine blade surface. This air provides a protective, insulating film which keeps the blade surface temperature well below the destructively high temperature levels of the combustor gases. This thesis presents for the first time, detailed pictures of the convective heat transfer distribution on the wall of an internal turbine blade passage near the entrance to a film cooling hole. The physical situation was modelled at 100X geometric scale as flow extraction into a single circular hole from a two-dimensional, fully developed, turbulent channel flow. High resolution heat transfer measurements were made using a transient technique with liquid crystals as surface temperature indicators. During the experiments, the two-dimensional channel Reynolds number was held constant while the flow extraction rate was varied for each of four hole inclination angles. The main region of heat transfer enhancement was found to be downstream with local heat transfer levels up to 6.5 times the levels associated with turbulent channel flow. Additional experimental, analytical, and computational flow field studies showed that the enhancement was caused mainly by the removal of the upstream boundary layer and the formation of a new laminar boundary layer at the down- stream hole edge. This new boundary layer was also influenced by downwash from a vortex pair.