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一、引言当导弹的头部为半球一圆柱组合体时,超音速气流流经组合体形成了很强的弓形冲波,使处于冲波之后翼面处的流场和远前方的流场有所差异。这时,已不能利用远前方的流场参数来计算导弹气动系数,需要确定半球—圆柱组合体头部后面的当地流场。在超音速情况下,求解半球—圆柱组合体冲波后的流场已有一些理论计算方法,如特征线法,但由于计算量较大,在工程中很难使用。在高超音速范围内,目前已研究出近似计算方法,比较简便,适合于工程使用。本文将此近似方法推广应用于超音速下绕
I. INTRODUCTION When the head of a missile is a hemisphere-cylinder combination, the supersonic airflow flows through the assembly to form a strong arcuate wave so that the flow field at the airfoil and the flow field far in front of Differences. At this time, it is no longer possible to calculate the aerodynamic coefficient of the missile using the far-field flow field parameters, and it is necessary to determine the local flow field behind the head of the hemisphere-cylinder combination. At supersonic speeds, there are some theoretical methods for solving the flow field of the hemispherical-cylinder assemblage after the wave, such as the characteristic line method, but it is difficult to use in the project due to the large amount of calculation. In the hypersonic range, approximate calculation methods have been developed so far, which are relatively simple and suitable for engineering use. In this paper, this approximate method is applied to supersonic winding