In modern gas turbine engines the combustor exit flow has a non-uniform temperature profile because of the discrete nature of the injection of fuel and dilution air, and the wall cooling flows. The affect of this non-uniform temperature profile on the aerodynamics and heat transfer rate of nozzle guide vanes and turbine blades is difficult to predict, and knowledge of this is important for estimating turbine component life and efficiency. Measurements of heat transfer have been conducted on an annular transonic intermediate pressure nozzle guide vane operating downstream of a high pressure rotating turbine stage. Measurements were made with and without a radial and circumferential inlet temperature profile. The experiments were conducted in the Isentropic Light Piston Facility (ILPF) at QinetiQ, a short duration engine size turbine facility with 1.5 turbine stages, in which Mach number, Reynolds number and gas-to-wall temperature ratios are correctly modelled. Experimental results are compared to predictions performed using boundary layer methods.

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