超声压气机叶栅流场的数值模拟与试验验证  

Numerical simulation and experimental investigation on supersonic compressor cascade flowfield

作　　者：唐凯[1,2] 葛宁[2] 顾杨[1] 向宏辉[1] 马昌友[1] 

机构地区：[1]中国燃气涡轮研究院,四川江油621700 [2]南京航空航天大学能源与动力学院,南京210016

出　　处：《燃气涡轮试验与研究》2015年第3期8-12,7,共6页Gas Turbine Experiment and Research

摘　　要：高超声压气机叶栅因适用于战斗机高马赫数飞行、增压比高而成为研究热点,但其损失难以控制,波系结构复杂,激波附面层干扰结果难以预测。基于自开发NUAA程序,对超声压气机平面叶栅流场进行计算分析,并通过与超声压气机平面叶栅试验结果的对比,考察叶栅在不同进口马赫数与气流攻角下的总性能、波系结构与激波位置。结果表明:程序计算的总性能与试验值吻合很好,且能精确捕捉超声叶栅中的激波结构,较好预测叶片表面等熵马赫数分布,可为超声叶栅的设计与结果验证提供支持。Supersonic compressor cascade has become a key issue for its high pressure ratio and applying to high speed fighter. But its loss is uncontrolled, the shock wave system is complex and the shock wave and boundary layer interaction are unpredicted. Based on NUAA program for supersonic compressor cascade flowfield, calculation analysis was conducted. And through the comparison between calculation results and experimental investigation, the performance parameters, shock formation and shock location at the condi-tion of different Mach number and attack angle were discussed. The results show that the NUAA program could correctly predict the shock wave structure and surface Mach number distribution, and the simulation results were good agreement with the experimental results. The program is helpful for supersonic compres-sor cascade design and testing result validation.

关 键 词：航空发动机 超声叶栅 激波 附面层 试验 数值计算 

分 类 号：V231.3[航空宇航科学与技术—航空宇航推进理论与工程]