STUDY OF FILM COOLING IN THE TRAILING EDGE REGION OF A TURBINE ROTOR BLADE IN HIGH SPEED FLOW USING PRESSURE SENSITIVE PAINT

摘要

This research focuses on film cooling of the trailing edge of a scaled up turbine rotor blade with engine-representative Mach number distribution. Pressure sensitive paint was used to obtain high-resolution adiabatic film cooling effectiveness measurements in the trailing edge region of the scaled turbine blade. The large scale, high-speed experimental set-up consists of a Perspex test section for maximum visibility of the PSP coated blade. The test section was designed to recreate a single blade passage of a gas turbine with inlet Mach and Reynolds numbers matching the corresponding values in an engine. The test blade has a constant cross section, representative of the mid-span profile of the high pressure turbine rotor blade. It was manufactured from aluminium to minimize temperature gradients over the surface of the test blade. In the current research, pressure surface cooling slots at the trailing edge were examined and the effect of cutback surface protuberance, or 'land', shapes on trailing edge film cooling was studied. Nitrogen and air were used as coolant gases giving a coolant to mainstream density ratio close to 1. Two land geometries-straight and tapered- were studied for a set of 6 blowing ratios from 0.4 to 1.4 in steps of 0.2. Land taper has a benefit for film cooling near the slot exit but its advantage reduces close to the trailing edge. For both geometries, film effectiveness falls with blowing ratio from 0.4 to 0.8 and increases with blowing ratio in the 0.8 to 1.4 range. Crossflow causes the coolant film to be biased towards one side of the lands. Film effectiveness results are compared with data from a scaled up low speed flat plat model of the trailing edge to explain the effect of acceleration on film cooling.