Abstract

Experimental data are provided for the surface of a blade tip within a transonic flow environment, which contains a squealer rim and a squealer recess. Four different film cooling arrangements are considered, each with holes placed at different locations along the upper pressure side of the turbine blade. The blade is mounted within a two-dimensional linear cascade, with four flow passages and five complete blades. Local, spatially-resolved and line-averaged distributions of surface adiabatic film cooling effectiveness and surface heat transfer coefficient ratios are provided for a range of local blowing ratios BR, with a ratio of tip gap to true blade span of 1.16 percent. Experimental and analytic procedures employed to obtain spatially-resolved film cooling surface data include infrared thermography, transient testing techniques, and the impulse response method for transient data analysis.

Results show that spatially-resolved surface thermal protection changes significantly as hole placement and flow condition of the upper pressure side film cooling are altered. Associated effectiveness values, along the upper pressure side of the turbine blade, vary in a significant manner as blowing ratio changes, with different characteristics depending upon upper pressure side film cooling hole placement. The influences of viscous dissipation within the turbine blade tip gap flow are quantified, which is vital to ascertain appropriate driving temperatures for convective heat transfer within such a high velocity, compressible environment. The influences of viscous dissipation effects are illustrated by adiabatic surface temperature magnitudes, measured with no film cooling, which, when considered relative to flow stagnation temperature, are directly related to local Mach number values within the tip gap flow along the blade tip surface.