The shear strain transport(SST) k-ω two-equation turbulence model was employed to numerically simulate the film cooling characteristics in the supersonic turbine cascade channel.The flow characteristics in the cascade channel and the change law of the film cooling efficiency were obtained at different gas film hole angles and blow ratios.The film jet near shock incident point can complement momentum to the boundary layer of separation zone

overcome the adverse pressure gradient and effectively improve the local overheating caused by shock wave.In subsonic flow state

the film incidence angle's impact on the efficiency of cooling can be embodied at larger blow ratio;on the contrary

in supersonic flow state

the film cooling efficiency is almost independent of the incidence angle

showing that the penetrating power of subsonic film to the supersonic mainstream is weaker than that to the subsonic mainstream;in the supersonic mainstream state

the film cooling efficiency of the shock incident location is higher than that of the downstream

which is mainly related to local turbulence intensity of the gas film.