Abstract: The effects of discrete film-cooling holes,wall curvature and the pressure gradient on the turbine blade film cooling are systematically studied by means of heat-mass analogy,numerical methods and flow visualization techniques.The flow and heat transfer performances of the cooling film over the flat plate,curved wall and blade surfaces are investigated.It is proven that the characteristics of discrete-hole film cooling are quite different from those of continuous-slot film cooling in the vicinity of the injection holes because the coolant jets from holes tend to separate from the cooled surfaces. An optimal blowing ratio is expected for ideal distribution of film cooling effectiveness.Wall curvature exerted an additional force on the cooling film.When momentum blowing ratio I(=ρ₂U²/ρ_∞U² _∞)>1,this force tends to seperate the film from the convex surface and to keep it close up to the concave surface,but when I<1,it acts contrary.For a given film cooling scheme,the effectiveness of turbine blade deponds on the blowing ratio,wall curvature and the pressure gradient.Some of the correlation formulas,exprimental data and computational results are presented to evaluate the film cooling effectivenesses and heat transfer coefficients at the leading edge,mid-chord and tailing edge of the turbine blade.