中文摘要：

为了研究超声速涡轮叶栅通道内的超声速气膜冷却,采用数值计算的方法,对主流压比2.33~4、冷气入射角度15°~45°条件下的涡轮叶栅超声速气膜流动和传热进行了研究.计算结果表明：超声速气膜射流与主流作用后产生的斜激波与尾缘激波交汇,形成两道反射激波,其中一道反射激波作用在气膜孔下游的叶片表面又形成了反射;在不同的主流压力下,超声速气膜射流在叶片法向和展向上展现出不同的发展特征,对转涡对（CVP）在展向上相互挤压,扼制了高温主流卷入叶片壁面;主流压比增加到4,气膜射流区在法向拉长,在展向相对较弱,导致主流在对转涡对（CVP）的作用下被卷入气膜射流的底层,壁面冷却效率降低;气膜入射角从15°增大到45°,冷却效率整体上呈先上升后下降趋势,在入射角30°时冷却效率相对最大,这与射流的穿透能力、冷却气流再覆壁面特征有关.

英文摘要：

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 de-veloping trend in the direction of the normal and span-wise of the vane surface under different mainstream pres-sure. The counter-rotating vortex pairs （CVP） extrude each other in span-wise aspect. It depresses the en-trainment 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-ro-tating 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 maxi-mum relatively. It has relationship with the reattachment of gas film and the penetrating capacity of the cooling jet.