An experimental investigation of turbine blade tip heat transfer and tip gap flows in the supersonic regime

Abstract

Gas turbine blade tip heat transfer and tip gap flow phenomena has been explored experimentally in a stationary cascade for blade exit Mach numbers = 1.2 to 1.4. Experimental results were found to agree well with qualitative predictions performed at GE Aircraft Engines.

The pressure distribution in the blade tip cavity of a grooved tip blade was found to vary little with either Mach number or tip gap height. The tip cavity pressure was, however, a strong function of location. The tip cavity pressure distribution coupled with the pressure side distribution near the tip was speculated to drive the leakage flow across the blade tip from mid-chord aft based on surface flow visualization studies using an oil/dye mixture.

Heat flux on the tip cavity floor was successfully measured using a thin-film Heat Flux Microsensor. Results of these measurements are consistent with previous studies in the subsonic regime. The convection coefficients on the tip cavity floor were found to be three times those found on the suction side airfoil surface near the trailing edge. Convection coefficients were found not to vary with either tip gap height or Mach number. The fluctuating component of heat flux was found to be at least 25% of the total heat flux.