Within a gas turbine engine, the turbine nozzle guide vanes are subjected to very harsh conditions from the highly turbulent and hot gases exiting the combustor. The temperature and pressure fields exiting combustors are highly nonuniform and dictate the heat transfer and aero losses that occur in the turbine passages. To better understand these effects, the goal of this work was to develop an adjustable combustor exit profile simulator for the Turbine Research Facility (TRF) at the Air Force Research Laboratory. The TRF is a high temperature, high pressure, short duration blow-down test facility that is capable of matching several aerodynamic and thermal nondimensional engine parameters including Reynolds number, Mach number, pressure ratio, corrected mass flow, gas to metal temperature ratio, and corrected speed. The primary research objective was to design, install, and verify a non-reacting simulator device that can provide representative combustor exit total pressure and temperature profiles to the inlet of the TRF turbine test section. This required the upstream section of the facility to be redesigned into multiple concentric annuli that serve the purpose of injecting high momentum dilution jets and low momentum film cooling jets into a central annular chamber, similar to a turbine engine combustor. The design of the section allows for variations in injection levels to generate different pressure profiles with elevated turbulence. The dilution and film cooling temperatures can also be varied to create a variety of exit temperature profiles similar to real combustors. The impact of the generated temperature and pressure profiles on turbine heat transfer and secondary flow development was ultimately investigated. Proposed optimal inlet conditions for the turbine tested in this research effort were determined based on the measured data corresponding to the combustor simulator exit profiles that minimized vane heat transfer and total pressure loss. / Ph. D.