基于膨胀波效应的高超声速进气道肩部流动分离控制研究

Investigation of flow separation controlling in hypersonic inlet shoulder based on expansion wave effect

为改善高超声速进气道唇口激波/附面层干扰诱导的肩部流动分离,从膨胀波及激波理论出发,推导出了膨胀波效应影响下的斜激波附面层干扰理论公式,获得了影响斜激波诱导分离的主要因素:膨胀角梯度、激波角及波前马赫数。在此基础上,开展了膨胀波效应影响下的流动分离控制研究,给出了膨胀波效应影响下斜激波诱导分离的判别及预测方法。结果表明:增大激波入射点处膨胀角梯度,可以显著减小甚至消除肩部流动分离;随着激波角增大,激波强度及逆压力梯度增加,分离区尺寸显著增大。而波前马赫数对分离区尺寸的影响不显著;在进口马赫数3.57~5.18,唇罩角度6°~10°范围内,当激波入射点处逆压比梯度小于250 m^(-1)时,斜激波诱导的流动分离消失,可为改善超声速/高超声速进气道内流道流动分离提供技术支撑。

In order to improve the flow separation induced by the cowl lip shock wave boundary layer inter⁃action in hypersonic inlet shoulder,based on the theory of expansion wave and shock wave,the theoretical formu⁃la of oblique shock wave boundary layer interaction under expansion wave effect was derived.And the main fac⁃tors of flow separation induced by shock wave boundary layer interaction were gained,which are the expansion angle gradient,the shock wave angle and the Mach number in front of shock wave.Based on this,the control of flow separation under the influence of expansion wave effect was studied,and the discrimination and prediction methods of oblique shock induced separation under the influence of expansion wave effect were given.The results show that increasing the expansion angle gradient at the shock wave incident point can significantly reduce or even eliminate flow separation of shoulder.With the increase of the shock wave angle,the shock intensity and the inverse pressure gradient increase,and the separation size increases significantly.However,the effect of the Mach number in front of shock wave on the size of separation is not significant.In the range of entrance Mach number from 3.57 to 5.18 and cowl lip angle from 6°to 10°,the flow separation induced by oblique shock wave disappears when the inverse pressure ratio gradient at the shock wave incidence point is less than 250 m^(-1),which can provide technical support for improving flow separation in supersonic/hypersonic inlet.