吸气式高超声速飞行器气动热试验研究  

Aeroheating Experiment for Airbreathing Hypersonic Vehicle

作　　者：徐大军[1] 蔡国飙[1] 乐川[1] 

机构地区：[1]北京航空航天大学宇航学院,北京100083

出　　处：《宇航学报》2006年第5期1004-1009,1095,共7页Journal of Astronautics

摘　　要：为获得吸气式高超声速飞行器气动热环境的数据,开展了气动热试验研究。在激波风洞中,来流马赫数Ma=6.12,来流单位雷诺数Re/L=1.37×107(1/m)试验条件下,对吸气式高超声速飞行器1/4缩比模型进行了表面气动热的测量。试验获得了小攻角变化范围内的飞行器头部前缘、头部上下交线、机身上下表面中心线、机身横截面周向、平尾垂尾前缘、发动机唇口等位置的热流率分布。研究结果表明,吸气式高超声速飞行器头部前缘、前体进气道壁面、发动机唇口、平尾垂尾前缘气动加热最为严重,另外乘波体外形的设计与布局影响热流的分布。To obtain aeroheating environment data of airbreathing hypersonic vehicle, aeroheating experiment had been conducted in shock wave wind tunnel. The test was carried at Maeh number 6.12 and unit Reynolds number Re/L = 1.37×10^7 （ 1/ m） with a 1/4 sub-scale model. The heat flux distribution at nose lead edge, nose upper/lower edge, airframe upper/lower centerline, body section outline, wing/fin lead edge were obtained. Results showed nose lead edge, airframe lower surface, inlet lip lead edge anti wing/fine suffered rigorous aeroheating environment, and configuration of waveraider influences heat flux distribution.

关 键 词：高超声速飞行器 乘波体 热流率 风洞试验 

分 类 号：O354.4[理学—流体力学] V211.73[理学—力学]