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

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.