Engine designers continually push the combustor exit temperature higher to produce more power from gas turbine engines. These high turbine inlet temperatures, coupled with high turbulence levels and flow field non-uniformities, make turbine vane and endwall cooling a very critical issue in engine design. To appropriately cool these surfaces, knowledge of the passage flow field and endwall temperature distribution at representative engine conditions is necessary. A combustor test section was used to simulate realistic turbine inlet profiles of turbulence, normalized temperature, normalized total pressure, and normalized streamwise velocity to study the flow field in a turbine vane passage and the adiabatic temperature distribution on the endwall. The combustor liner film-cooling and exit slot flows were varied independently to determine their relative effect on endwall cooling in the downstream turbine vane. Flow field measurements revealed the presence of a previously unmeasured third vortex in the vane passage. The tertiary vortex was located above the passage vortex and had rotation opposite to the passage vortex. Increasing the amount of slot flow reduced the size and strength of the nearwall vortices, while increasing the size and strength of the tertiary vortex. Adiabatic endwall temperature measurements revealed higher temperatures surrounding the base of the vane. The endwall measurements also showed that the exit slot flow was effective at cooling only a region of the endwall near the vane leading edge on the suction side. Increasing slot flow was found to have a larger thermal benefit to the endwall relative to increasing combustor liner film-cooling. / Master of Science