Systematic isothermal investigations on the aerodynamic effects of leading edge film cooling were carried out on a large scale high pressure turbine cascade named AGTB. In the vicinity of the stagnation point the AGTB turbine cascade has one injection site on the suction side and one on the pressure side. Three injection geometries were tested: Slots (two dimensional geometry), streamwise inclined holes (symmetrical three dimensional geometry) and compound angle holes (fully three dimensional geometry). The injection angle in streamwise direction, the blowing ratio, the inlet turbulence intensity, the inlet Mach number, and the inlet Reynolds number were kept constant at values typically found in modern gas turbines.
The measured data comprise the coolant plenum state, the cascade inlet conditions, the flow field in the cascade exit plane including secondary flows, the static pressure distribution in the mid span section of the blade and in the near hole region, the coolant flow field close to the injection site on the leading edge, Schlieren images of the coolant penetration height and oil-and-dye flow visualizations of the blade surface. The experimental data are summarized and documented as a test case that can be used for validation purposes of prediction methods.
