跨音涡轮叶栅吸力面气膜/激波交互作用的数值研究  

Numerical Study of Film / Shock Interaction on Suction Surface of Transonic Turbine Cascade

作　　者：黄康 陈帝云 李世峰 时培杰 马护生 Kang Huang;Di-yun Chen;Shi-feng Li;Pei-jie Shi;Hu-sheng Ma(Aerospace Technology Institute,China Aerodynamics Research and Development Center)

机构地区：[1]Aerospace Technology Institute,China Aerodynamics Research and Development Center

出　　处：《风机技术》2022年第6期40-47,共8页Chinese Journal of Turbomachinery

基　　金：National Natural Science Foundation of China(51806233);National Major Science and Technology Project(2017-III-0003-0027)。

摘　　要：The interaction between the cooling film on the suction surface of the transonic turbine cascade and the shock wave in the passage is analyzed by numerical calculation method.The effects of cooling pressure and incident angle on the shock wave shape and the film cooling efficiency under the action of shock wave are studied.The results show that with the increase of exit Mach number,the incident position of the left shock wave at the trailing edge of the cascade moves backward relative to the leading edge of the adjacent cascade,and the shock wave intensity increases;there is a temperature jump behind the temperature peak at the exit Mach number of 1.05 and 1.2,the main reason is that with the increase of shock wave intensity,the interaction between shock wave and boundary layer intensifies,and the boundary layer appears local separation,in the separation region,the mixing of high-temperature mainstream and boundary layer flow intensifies,which makes the temperature in the separation region of suction surface increase;the results show that different cooling gas injection positions have great influence on the shock wave shape in the channel,and the cooling area of the wall is smaller than that before and after the shock wave,which is mainly due to the reduction of the film coverage area under the impact of the incident shock wave,after the impact zone of shock wave,the cooling film is compressed by the main stream and attached to the wall again,resulting in the discontinuous film coverage at the shock wave incident point;the film cooling efficiency decreases significantly with the increase of cooling flow angle from 25°to 35°and changes little with the increase of cooling flow angle from 35°to 45°;the cooling efficiency does not increase infinitely with the increase of blowing ratio,when the blowing ratio is greater than a certain value,the influence of blowing ratio on cooling efficiency gradually decreases.

关 键 词：Transonic Turbine Suction Surface Film Cooling Shock Wave 

分 类 号：V231[航空宇航科学与技术—航空宇航推进理论与工程]