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采用标准k-ωSST湍流模型数值计算方法,针对二元高超声速倒置进气道激波与附面层严重的干扰现象,采用分流楔抑制激波与附面层干扰方法,对有无分流楔的进气道性能及流动机理特征进行了详细的研究。结果表明:采用分流楔的流动控制方法,有效抑制了激波/附面层产生的分离包对进气道内流动的干扰;提高倒置进气道的气动性能,进气道的总压恢复系数和流量捕获系数均有提高,计算模型的壁面总阻力系数得到一定程度的减小。数值计算结果表明,在分流楔尾迹中强剪切流动在一定程度上缓解了激波与附面层干扰的强度;在分流楔后缘存在稳定的横向涡,由于气流进入尾迹驻涡是来自附面层外的总压较高的高能流体,提高了附面层的抗逆压能力;由于尾迹驻涡的存在使得分离涡没有向弹体周向扩散,减小了阻力。该方法实现了对高超声速倒置进气道激波/附面层干扰的抑制,揭示了其抑制的机理。
A standard k-ωSTST turbulence model is used to calculate the numerical value of the turbulence model. In order to deal with the serious interference phenomenon between the shock wave of baryonic supersonic inlet and the surface layer, a shunt wedge is used to suppress the interference between the shock wave and the adherent layer. Inlet characteristics and flow characteristics of a detailed study. The results show that the shunt wedge flow control method can effectively suppress the interference of the separation bag generated by the shock wave / surface layer on the flow in the inlet duct and improve the aerodynamic performance of the inverted inlet duct, the total pressure recovery coefficient of the inlet duct and The flow capture coefficient all increased, and the total wall drag coefficient of the model was reduced to a certain degree. The results of numerical calculation show that the strong shear flow in the wedge wake of the shunt dissipates the intensity of the interference between the shock and the surfacing layer to a certain degree. There is a stable transverse vortex at the trailing edge of the shunt wedge. Because of the presence of wake in the vortex, the separation vortex does not diffuse to the perimeter of the projectile, reducing the resistance. The proposed method suppresses the shock / overcoat interference of the hypersonic inverted inlet, and reveals the mechanism of its suppression.