高超声速进气道前缘钝度效应试验研究  

Experimental Studies of Leading Edge Bluntness Effects on Hypersonic Inlet

作　　者：张红军[1] 沈清[1] 

机构地区：[1]中国航天空气动力技术研究院,北京100074

出　　处：《推进技术》2013年第10期1316-1320,共5页Journal of Propulsion Technology

基　　金：航天技术自主研发基金;国家自然科学基金(90816026)

摘　　要：基于一种典型高超声速二元进气道,考察前缘钝度效应对进气道边界层转捩的影响,加工了四种半径为R=0.05mm,R=0.1mm,R=0.2mm,R=0.25mm的前缘,在FD-07风洞中开展了自然转捩及人工转捩的风洞试验。试验中采用压缩拐角压力分布特征及进气道起动相结合的方法来估计边界层转捩位置,得出了进气道压缩面边界层转捩位置随前缘半径变化的规律。试验表明在来流条件下随前缘钝化半径增加,边界层转捩位置明显后移。针对R=0.25mm时进气道不起动的情况,基于线性稳定性理论(LST)理论设计了人工转捩条带,通过试验成功实现了转捩。The leading edge bluntness effects on inlet boundary layer transition were studied based on a typical two dimensional hypersonic inlet. Four models of different leading edge radius （R = 0.05mm, R = 0.1 mm, R =0.2mm, R = 0.25ram） were studied in FD-07 wind tunnel, including natural transition and ar- tificial transition. The location of boundary layer transition was identified through the corner pressure distri- bution characteristics and inlet starting, the rule of boundary layer transition position with the variation of leading edge radius was obtained. Results show that, under the wind tunnel condition, the boundary layer transition position moves down stream with increasing the leading edge radius. The inlet will not start when the leading edge radius R=0.25mm. By designing artificial transition strips based on linear stability theory （LST） theory, the inlet is started successfully.

关 键 词：高超声速进气道 钝度效应 边界层转捩 

分 类 号：V435.12[航空宇航科学与技术—航空宇航推进理论与工程]