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为了在级间分离期间提供反推力，许多固体火箭发动机前端都装有一组斜切反喷管。由于反喷管的气动型面具有许多尖点，并且在超声速区有一个台阶，喷管内存在一系列激波，而且亚声速区和超声速区互相混杂。本文用时间相关法数值模拟了反喷管流场；控制方程用ＭａｏＣｏｒｍａｃｋ显格式求数值解；边界参数采用物理边界条件和从双特征线方法推导来的有效的特征方程来计算；对于固壁边界点，在不同的区域采用不同的方法对计算方程组求数值解；对超声速台阶区，采用Ｃｈａｐｍａｎ—Ｋｏｒｓｔ模型计算。计算得到的壁面压强分布与风洞吹风实验测得的压强分布一致，计算得到的等密度线分布规律与纹影照片上的密度变化规律一致。
In order to provide thrust reversal during interstage separation, many solid rocket motors are equipped with a series of chamfered counter-flow nozzles on the front end. Because the aerodynamic profile of the nozzle has many sharp points and there is a step in the supersonic zone, there is a series of shocks in the lance and the subsonic and supersonic zones are intermingled. In this paper, the numerical simulation of the nozzle flow field with time-dependent method; the governing equation using MaoCormack explicit format for numerical solution; boundary parameters using physical boundary conditions and the effective characteristic equation derived from the double characteristic line method to calculate; for the solid wall boundary In different regions, different methods are used to solve the numerical equations. For the supersonic steps, the Chapman-Korst model is used. The calculated wall pressure distribution is consistent with the pressure distribution measured by wind tunnel blowing experiment. The distribution of isosceles line is consistent with the change of density on the smear photograph.