抄録
This paper deal with the film cooling effectiveness of round and shaped hole on the suction side of the airfoil under transonic condition. The measurement with heat - mass transfer analogy has been conducted at linear turbine cascade facility to investigate the effect on film cooling effectiveness due to varying exit Mach number and blowing ratio. Shock boundary layer interaction scarcely has influence on the film cooling effectiveness, and film flow is hard to diffuse in the lateral direction under transonic condition. In addition, the comparison between these experimental data and the results with CFD (Computational Fluid Dynamics) is reported.
