基于后掠唇罩的入射激波/边界层干扰流动控制方法

Flow control method for impinging shock/boundary-layer interaction based on swept-back cowl configuration

针对高超声速进气道内强唇罩激波/边界层干扰带来的大尺度分离、流动损失大等问题,提出了基于后掠唇罩的入射激波/边界层干扰流动控制方法。在来流马赫数3、唇罩压缩角18°条件下,仿真对比了后掠/平直2类唇罩2种构型干扰区内的流动特性。结果表明:后掠唇罩入射激波/边界层干扰产生的分离包尺度沿展向呈现逐渐增加的趋势,利用三维后掠入射激波产生的自对称面往两侧的顺压梯度,驱动低能流往两侧迁移,可使得分离区流向尺度相对于平直唇罩构型最大可减小50.6%;在后掠唇罩干扰区内,压力分布呈现以分离线曲率中心为虚拟中心的椭圆相似特征。

The strong cowl shock wave/boundary-layer interaction within the hypersonic inlet induces large scale sepa⁃ration and strong energy loss.To address this problem,we propose a flow control method for impinging shock wave/boundary-layer interaction based on the swept-back cowl configuration.The flow characteristics with swept/unswept cowl configurations were numerically revealed with the incoming Mach number of 3.0 and the cowl compression angle of 18°.The results show that the length-scale of the separation,induced by the swept-back cowl shock wave/boundary-layer interaction,increases along the spanwise direction.By utilizing the lateral pressure gradient resulted from three-dimensional swept shock waves,it forces the low momentum flow to migrate from the symmetry plane to the sidewall,leading to an over 50.6%diminution of the separation bubble compared to that of the unswept cowl con⁃figuration.In addition,the pressure distribution in the interaction region shows an elliptical similarity with the center of separation line curvature as the virtual center.