ABSTRACT The design and development of current and future gas turbine engines for aircraft propulsion have focused on operating the high pressure turbine at increasingly elevated temperatures and pressures. The drive towards thermal operating conditions near theoretical stoichiometric limits as well as increasingly stringent requirements on reducing harmful emissions, both equate to the temperature profiles exiting combustors and entering turbines becoming less peaked than in the past. This drive has placed emphasis on determining how different types of inlet temperature and pressure profiles affect the first stage airfoil endwalls. The goal of the current study was to investigate how different radial profiles of temperature and pressure affect the heat transfer along the vane endwall in a high pressure turbine. Testing was performed in the Turbine Research Facility located at the Air Force Research Laboratory using an inlet profile generator. Results indicate that the convection heat transfer coefficients are influenced by both the inlet pressure profile shape and the location along the endwall. The heat transfer driving temperature for inlet profiles that are nonuniform in temperature is also discussed. INTRODUCTION The performance and durability of the hot section within gas turbine engines are critical operational issues that present many design and research challenges. The hot section of these engines includes both the combustion chamber and the high pressure turbine, the latter of which includes the endwall regions under investigation in this study. Considering that the hot gas temperatures are well above the melting point of the metal turbine hardware, the heat transfer to and aerodynami
