非均匀流等压比变后掠角高超侧压式进气道研究

INVESTIGATION OF SIDEWALL COMPRESSION INLET WITH CONSTANT PRESSURE RATION AND CURVED LEADING EDGE UNDER NONUNIFORM SUPERSONIC FLOW

关键词:  非均匀流  进气道试验  高超声速进气道  风洞试验

Key words:  Nonuniform flow  Inlet test  Hypersonic inlet  Wind tunnel test

摘要:

通过理论分析和风洞实验，对工作在前体附面层内的侧压式进气道，研究了等激波压比和等溢流角前提下侧压缩面的设计方法，分析了６种不同的侧压缩型面在４种来流附面层中，波后压力沿高度的变化规律和溢流角的变化规律。研究发现，采用部分圆弧加直线为前缘。四次曲线为斜面后缘型线的侧压缩面，在４种非均匀来流下的特性较好。马赫５．３的非均匀流风洞实验结果表明，等压比和等溢流角设计的侧压式进气道较通常的直前缘侧压式进气道，在非均匀来流中喉道截面马赫数分布均匀度好，总压恢复略高

Abstract:

Three dimensional sidewall compression inlet model with constant pressure ratio p 21 and constant spillage angle for nonuniform supersonic incoming flow was designed and experimentally investigated.The design of compression angle and leading edge sweep was based upon the criterion of a constant pressure ratio and constant spillage angle when a typical boundary layer flow was swallowed by the inlet.Three inlet models were tested in Ma =5 3 wind tunnel.The model 1 with leading edge sweep of 30° acted as a baseline inlet.The new design,named model 3,with a partly curved leading edge,partly constant sweep leading edge and 4th power curved trailing edge was specially designed for nonuniform incoming flow.The experimental results indicate that the model 3 performs better than the baseline configuration model 1 in terms of Mach number in throat and total pressure recovery