一种超声速进气道侧壁面流场控制研究

A Side Wall Flow-Control Method of Supersonic Inlet

关键词:  超声速进气道  流动控制  进气道性能  激波  边界层干扰

Key words:  Supersonic inlet  Flow control  Inlet performance  Shock wave  Boundary interaction

摘要:

为改善二元超声速进气道前体激波与侧壁面边界层干扰问题，提出了一种在侧壁开泄流气缝的流场控制方法并进行了数值仿真验证，然后研究了侧壁面开缝的宽度、位置、角度等典型几何参数对进气道性能的影响规律。结果表明：设计马赫数下侧壁开缝使进气道唇口角区处的溢流明显减小，进气道内通道进口流场得到改善，进气道流量系数提高2.27%，喉道截面总压恢复系数提高3.37%；在非设计状态下，进气道性能也有一定的改善。典型几何参数研究结果表明，当侧壁开缝位置位于前体斜激波位置(L=-1.4~-0.21)、开缝宽度为0.85~1.10倍当地边界层厚度时，对进气道性能的改善效果最佳，而开缝的角度影响并不明显。

Abstract:

To improve the interaction between forebody shock wave and side wall boundary layer of 2-D supersonic inlet， a flow-control method of bleeding on side wall is put forward. The influence rules of typical geometric parameters such as width， position and angle of side wall slit on the performance of the inlet are studied by numerical simulation. The results show that bleeding on side wall decrease spillage of lip corner and improve inlet channel flow field under design Mach number condition. The inlet flow ratio increased by 2.27%， and the inlet throat total pressure recovery coefficient increased by 3.37%. In addition， the performance of inlet is also improved under off-design condition. The research results of typical geometric parameters indicate that performance of the inlet have been improved obviously when side wall slit is located at position of forebody shock wave (L=-1.4~-0.21) and slit width equals 0.85~1.10 times thickness of boundary layer. The effect on the range of slit angle researched can be ignored.