摘要
本文对音速及等压两种不同的出口边界下跨音速扩压器的流场做了实验研究.扩压器的扩张角为6°,激波波前马赫数均为1.335.实验结果表明:在等压出口边界条件时,激波自恃振荡的幅度及分离包的尺度比音速出口条件时要大得多.
Experimental investigation of normal shock wave/turbulent boundary layer interaction flow field in a transonic diffuser was made between the choked and the non-choked exit conditions. The divergent angle of the diffuser is 6 Deg. The Mach numbers before the shock wave were 1.335, The experimental results showed that, in the non-choked exit-boundary condition, the RMS of shock oscillation and the scale of the separation bubble as well as the boundary-layer growth are greater than those in the choked. For the straight-wall diffuser, the ratio of RMS of shock oscillation to the length of separation bubble is near constant.
出处
《推进技术》
EI
CAS
CSCD
北大核心
1991年第1期19-23,共5页
Journal of Propulsion Technology
关键词
扩压器
超音速
进气道
流场
非定常流动
飞机
Diffuser, Supersonic inlet, Boundary condition, Flow field, Unsteady flow
