Abstract
The effects of upstream injection angle on film cooling effectiveness of a turbine vane end wall with various endwall film-hole designs were examined by applying pressure-sensitive paint (PSP) measurement technique. As the leakage flow from the slot between the combustor and the turbine vane is not considered an active source to protect the vane endwall in certain engine designs, discrete cylindrical holes are implemented near the slot to create an additional controllable upstream film to cool the vane end wall. Three potential injection angles were studied: 30 deg, 40 deg, and 50 deg. To explore the optimum endwall cooling design, five different film-hole patterns were tested: axial row, cross row, cluster, midchord row, and downstream row. Experiments were conducted in a four-passage linear cascade facility in a blowdown wind tunnel at the exit isentropic Mach number of 0.5 corresponding to inlet Reynolds number of 380,000 based on turbine vane axial chord length. A freestream turbulence intensity of 19% with an integral length scale of 1.7 cm was generated at the cascade inlet plane. Detailed film cooling effectiveness for each design was analyzed and compared at the design operation conditions (coolant mass flow ratio (MFR) 1% and density ratio 1.5). The results are presented in terms of high-fidelity film effectiveness contours and laterally (spanwise) averaged effectiveness. This paper will provide the gas turbine designers valuable information on how to select the best endwall cooling pattern with minimum cooling air consumption over a range of upstream injection angle.
