Transonic Film Cooling Effectiveness from Shaped Holes on a Simulated Turbine Airfoil

摘要

The performance of film cooling on a simulated transonic, turbine airfoil is investigated using an experimental apparatus designed especially for this purpose. A symmetric airfoil with the same transonic Mach number distribution on both sides is employed. Mach numbers along the airfoil surface range from 0.4 to 1.24 and match values on the suction surface of airfoils from operating aeroengines. Film cooling holes are located on one side of the airfoil near the passage throat where the freestream Mach number is nominally 1.07. Three different film cooling configurations are investigated with density ratios of about 1.4-1.6 over a range of blowing ratios, momentum flux ratios, and pressure ratios. These include single rows of 1) cylindrical holes with simple angle orientations, 2) laid-back-fan-shaped holes with simple angle orientations, and 3) laid-back-fan-shaped holes with compound angle orientations. New spatially resolved and spatially averaged effectiveness data and discharge coefficients are provided for these hole geometries with airfoil suction surface geometry, flow conditions, and boundary-layer development that simulate engine operating conditions. The highest local and spatially averaged film cooling effectiveness magnitudes are generally obtained with laid-back-fan-shaped holes oriented with compound angles, followed by laid-back-fan-shaped holes oriented with simple angles, which are then followed by round, cylindrical holes oriented with simple angle orientations.