Impact of Turbine Inlet Temperature Distortion on Rotor Blade Heat Transfer at Engine Representative Conditions Using Infrared Thermography

Abstract

With inlet gas temperatures of high-pressure turbine stages of engines reaching almost 2000 K, shroudless turbine blade tips have become a performance and life critical component. As temperature distortions have a direct impact on rotor blades heat transfer, there is a need to gain experimental insights with highly engine representative conditions. This paper presents heat transfer measurements, scaleable to engine conditions, from a rotating transonic turbine test facility using a novel infrared measurement technique [1]. Experiments were conducted in the Oxford Turbine Research Facility (OTRF), a rotating facility capable of matching engine Mach and Reynolds numbers, non-dimensional speed and gas-to-wall temperature ratio. The work investigates the effect of a temperature profile typical of a rich burn combustor in a modern civil engine on the heat transfer in the high-pressure turbine rotor. Two inlet conditions are investigated: (1) uniform temperature and pressure distribution, and (2) radial temperature and uniform pressure distribution. The single stage high-pressure turbine includes cooled nozzle guide vanes and uncooled rotor blades with a squealer tip geometry. High-resolution two-dimensional transient measurements of surface temperature are conducted using infrared thermography, processed to heat flux and used to determine adiabatic wall temperature and Nusselt number for the blade tip and pressure side, a result not achieved before. Comparison to thin-film gauge heat transfer measurements on the blade pressure surface highlight the capabilities of the infrared thermography system. A computational simulation conducted using the Rolls-Royce code HYDRA is used to demonstrate the confidence in the novel experimental measurements.