大攻角范围超声压气机叶栅激波波系研究  

Investigation on the shock wave system of supersonic compressor cascade at different attack angle

作　　者：唐凯[1,2] 葛宁[2] 顾杨[1] 向宏辉[1] 温珍荣[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》2016年第2期12-15,20,共5页Gas Turbine Experiment and Research

摘　　要：对超声压气机叶栅的多攻角工况进行试验,利用纹影仪、油流试验及叶片表面等熵马赫数分布结果进行对比分析,观察到大攻角范围下叶栅激波波系结构发生了明显变化。为揭示激波结构变化原因,利用NUAA计算程序对叶栅进行仿真。研究发现,大攻角状态下叶栅通道中斜激波产生的原因,为前通道激波诱发附面层分离再附后,气流为沿叶片表面继续流动,从而形成斜激波;由于斜激波的增压降速,导致尾缘激波非常微弱甚至消失。An experiment on supersonic compressor cascade at different attack angle was carried out. By comparing with the results on the schlieren apparatus, oil flow visualization and surface Mach number distri-bution, a conclusion was made that the structure of the shock wave system changed a lot at high attack an-gle. To explain the reasons, CFD based on NUAA program was used. The results revealed that the oblique shock wave was caused when boundary layer separated and reattached, the air flowed along the blade sur-face. The trailing edge shock wave got weak or even disappeared after the velocity decreased by the oblique shock wave.

关 键 词：压气机 超声叶珊 激波 波系结构 附面层 大攻角 试验 数值仿真 

分 类 号：V232.4[航空宇航科学与技术—航空宇航推进理论与工程]