The effects of suction surface film cooling on aerodynamic losses are 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 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 free-stream Mach number is nominally 1.07. Round cylindrical and conical diffused film cooling hole configurations are investigated with density ratios from 0.8 to 1.3 over a range of blowing ratios, momentum flux ratios, and Mach number ratios. Also included are discharge coefficients, local and integrated total pressure losses, downstream kinetic energy distributions, Mach number profiles, and a correlation for integral aerodynamic losses as they depend upon film cooling parameters. The contributions of mixing and shock waves to total pressure losses are separated and quantified. These results show that losses due to shock waves vary with blowing ratio as shock wave strength changes. Aerodynamic loss magnitudes due to mixing vary significantly with film cooling hole geometry, blowing ratio, Mach number ratio, and (in some situations) density ratio. Integrated mixing losses from round cylindrical holes are three times higher than from conical diffused holes, when compared at the same blowing ratio. Such differences depend upon mixing losses just downstream of the airfoil, as well as turbulent diffusion of streamwise momentum normal to the airfoil symmetry plane. [S0889-504X(00)02202-9]

The effects of suction surface film cooling on aerodynamic losses are 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. Round cylindrical, and conical diffused film cooling hole configurations are investigated with density ratios from 0.8 to 1.3 over a range of blowing ratios, momentum flux ratios, and Mach number ratios. Also included are discharge coefficients, local and integrated total pressure losses, downstream kinetic energy distributions, Mach number profiles, and n correlation for integral aerodynamic losses as they depend upon film cooling parameters. The contributions of mixing and shock waves to total pressure losses are separated and quantified. These results show that losses due to shock waves vary with blowing ratio as shock wave strength changes. Aerodynamic loss magnitudes due to mixing vary significantly with film cooling hole geometry, blowing ratio, Mach number ratio, and (in some situations) density ratio. Integrated mixing losses from round cylindrical boles are three times higher than from conical diffused holes, when compared at the same blowing ratio. Such differences depend upon mixing losses just downstream of the airfoil as well as turbulent diffusion of streamwise momentum normal to the airfoil symmetry plane.