Abstract
A wind tunnel test section was designed to simulate the effects of streamwise pressure gradient on the boundary layer developing on the suction surface of a gas turbine airfoil. This study was part of an ongoing research program investigating various aspects of turbine film cooling and is the groundwork for future studies involving film cooling effectiveness in the presence of mainstream pressure gradients. The pressure distribution under consideration was typical of those encountered on airfoils in high-pressure, high-work turbines that are currently being developed for high-performance aircraft applications. Numerical simulations were performed using the 2-D boundary layer code TEXSTAN to predict the accelerating flow field in the turbine based on sponsorsupplied airfoil data. By matching the nondimensional acceleration and Reynolds numbers from the numerical simulation, a contoured roof was developed for the test section of an existing wind tunnel that would induce a flow at low speed and low temperature that is dynamically similar to the flow across the turbine blade. Experimental verification of the facility was accomplished with wall static pressure measurements and LDV velocity measurements.
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