关键词: 超声速涡轮叶栅；超声速气膜冷却；数值模拟；流动特征；冷却特性

Abstract: In order to study supersonic film cooling on the supersonic turbine cascade，some numerical simulations were carried out to present the flow of gas film and heat transfer characteristics on condition of the pressure ratio of 2.33 ~ 4，the injecting angle of 15 ° ~ 45 °. The results show that oblique shock wave deduced from the interaction between supersonic gas film and mainstream intersects with the trailing edge shock. Then，two reflected shocks come into being. One of the two reflected shocks acts on the turbine vane surface located on the downstream of the film hole to form the refection of shock again. Supersonic gas film exhibits the different developing trend in the direction of the normal and span-wise of the vane surface under different mainstream pressure. The counter-rotating vortex pairs (CVP) extrude each other in span-wise aspect. It depresses the entrainment of high temperature mainstream into the blade wall. When the mainstream pressure reaches to 4，the area dominated by the gas film stretches longer in normal and develops relatively weaker in span-wise. It results in the decrease of cooling effectiveness for the mainstream involved into the bottom of gas film by the counter-rotating vortex pairs (CVP). When the angle of the film hole increases from 15 ° to 45 °，the cooling efficiency increases firstly and then decreases. While the incidence angle is 30 °，the cooling efficiency reaches to maximum relatively. It has relationship with the reattachment of gas film and the penetrating capacity of the cooling jet.

Key words: Supersonic turbine cascade；Supersonic film cooling；Numerical simulation；Flow characteristics；Cooling characteristics