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文摘
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在JF-8脉冲风洞中,来流马赫数Ma＝8.0,来流单位长度雷诺数Re/L=1.47×107和2.52×107(1/m)两种试验条件下,对高超声速飞行器1/20缩尺模型进行了表面气动热的测量.模型攻角α=0°,10°,15°,20°,25°和30°.试验给出机身对称面、翼前缘、立尾前缘等处的热流率分布.机头部分最大热流率与由Fay-Riddell公式计算的驻点热流Q0率接近,翼前缘最大热流率在全机身中最大,约为Q0的2倍,因此翼前缘的热环境是最严酷的. |
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其他语种文摘
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The aerodynamic heat flux of the hypersonic vehicle model is measured in JF-8 impulsive wind tunnel. The test is carried out at Mach number 8.0 and two unit Reynolds number Re/L = 1.47 * 10~7 and 2.52 * 10~7(l/m) . The attack angle range is from 0 to 30 degrees. The heat flux distribution is provided at the asymmetric side of the body, wing and the vertical fin. The maximal heat flux at the nose is very similar with that at the stagnation point. The maximal heat flux of the whole vehicle is at the wing leading edge, which is approximately 2 times of the value at stagnation point. So the most rigorous thermal environment is near the wing leading edge. |
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来源
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流体力学实验与测量
,2004,18(1):29-32,37 【核心库】
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关键词
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高超声速飞行器
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热流率
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风洞试验
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地址
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中国科学院力学研究所LHD, 北京, 100080
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语种
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中文 |
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文献类型
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研究性论文 |
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ISSN
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1007-3124 |
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学科
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航空 |
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文献收藏号
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CSCD:1500723
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