一、前言 跨超声速压气机的研制与发展,促进了对超声速来流叶栅的研究。当来流的轴向速度分量为亚声速时,存在着唯一进气角关系式β_∞=f(M_∞)。在准确计算这一数值关系时,就会碰到一个如何确定栅前激波形状和位置的问题。对于单翼或孤立钝体已有了若干近似方法,用来确定脱体激波的形状和位置。但对于叶栅,由于激波前的流场是不均匀的,与单翼的情况不同,至今还没有一个叶栅的激波模型。因此,世界各国的研究者 I. INTRODUCTION The development and development of transonic supersonic compressor has promoted the study of supersonic flow cascades. When the axial velocity component of subsurface flow is sub-sonic, there exists the only relation of inlet angle β_∞ = f (M_∞). Accurately calculate the numerical relationship, you will encounter a how to determine the shape and location of pre-grid shock wave problem. There are several approximate methods for single-wing or isolated bluff bodies used to determine the shape and location of the off-body shock. However, for the cascade, there is not a cascade shock model due to the non-uniform flow field before the shock wave. Therefore, researchers from all over the world