Abstract: To study the effect of leading edge radius and model scale on flowfield structure, two scale models of three-stage ramp was used. The length of the test models is 0.3m and 0.6m, respectively, and the leading edge radius is 0mm, 0.5mm, 1mm, 1.5mm and 3.0mm, respectively. The experiment was carried out in 0.6m Shock Tunnel of CARDC, and the test Mach number is 5.98with total temperature of 670K and total pressure of 6.557MPa. The shadowgraph photos series were obtained using a high speed camera. The shock wave angle data from the photos indicate that the first shock wave angle increases, while the second and third decrease as the leading edge radius increasing. The shock wave angles of the two scale test models are obviously different, which means scale effect exists indeed on the flowfield structure. For the two scale models with the same leading edge bluntness (except the sharp leading edge model), the difference of the first shock wave angle is about 0.4°, and the maximum difference between the second and third shock wave angle is about 0.5°. The flowfield visualization also shows that the second and third shock wave roots are curved. Shock wave boundary layer interaction regions at the compression corners are clearly observed. It is also observed that the shock wave roots are curved more seriously and its influence area extends with the leading edge bluntness increasing. 

Key words: Multistage compression ramp; Hypersonic inlet; Leading edge radius; Shock wave boundary layer interaction; Shock wave angle

摘要： 为研究前缘钝度及模型尺度对流场结构的影响,采用了长度为0.3m和0.6m的三级压缩楔模型,前缘半径分别为0,0.5,1,1.5,3mm,在0.6m激波风洞中利用高速阴影摄像获得了系列流场结构照片,清晰地显示了激波结构。试验条件为马赫数5.98,总温670K,总压6.56MPa。数据结果表明,随着前缘半径的增加,第一道激波角增大,第二和第三道激波角减小；存在明显的模型尺度影响,在同等钝度条件下(尖前缘除外),两个尺度模型的第一道激波角相差达0.4°,第二道和第三道激波角最大可相差0.5°。流场照片显示,在拐角处存在激波边界层干扰,造成第二、三道激波根部弯曲,随前缘半径增加,弯曲程度和影响区域增大。 

关键词: 多级压缩楔;高超进气道;前缘半径;激波边界层干扰;激波角 