Experimental Investigation of the Effect of Engine Representative Combustor Exit Temperature Profile and Disc Cavity Leakage Flow on the Film Cooling of Contoured Hub Endwall of a High Pressure Gas Turbine Rotor Cascade

The effects of an engine representative combustor exit temperature profile on the adiabatic effectiveness values on a contoured turbine hub endwall at different disc cavity leakage flow rates are documented. Measurements for passage thermal fields and endwall adiabatic effectiveness values are made for disc cavity leakage mass flow rates of 0.5%, 1.0 % and 1.5 % of the mainstream passage flow; in presence of an engine representative combustor exit temperature profile. The endwall contour resembles the shape of a 'shark nose'. Endwall adiabatic effectiveness values for the 'shark nose' profile have been compared with those with a different endwall profile called 'dolphin nose'. Also, augmentation in adiabatic effectiveness by using an engine representative combustor exit temperature profile relative to other profiles has been documented. The results indicate that increasing leakage flow rates leads to a more uniform distribution of adiabatic effectiveness values in the pitchwise direction at the passage entrance. Also, the presence of additional coolant in the near wall approach flow, leads to significant endwall cooling near the passage exit.