Transonic aerodynamic losses due to turbine airfoil, suction surface film cooling

The effects of suction surface film cooling on aerodynamic losses are inves tigated using an experimental apparatus designed especially for this purpos e, 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 march value s on the suction surface of airfoils from operating aeroengines. Film cooli ng holes are located on one side of the airfoil near the passage throat whe re the free-stream Mach number is nominally 1.07. Round cylindrical and con ical diffused film cooling hole configurations are investigated with densit y ratios from 0.8 to 1.3 over a range of blowing ratios, momentum flux rati os, and Mach number ratios. Also included are discharge coefficients, local and integrated total pressure losses, down stream kinetic energy distribut ions, Mach number profiles, and a correlation for integral aerodynamic loss es as they depend upon film cooling parameters. The contributions of mixing and shock waves ro total pressure losses are separated and quantified. The se results show that losses due to shock waves vary with blowing ratio as s hock wave strength changes. Aerodynamic loss magnitudes due to mixing vary significantly with film cooling hole geometry, blowing ratio, Mach number r atio, 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 upo n mixing losses just downstream of the airfoil, as well as turbulent diffus ion of streamwise momentum normal to the airfoil symmetry plane.