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为了研究壁温对高超声速飞行器阻力的影响,在常规高超声速风洞和脉冲燃烧加热风洞中开展试验研究,结合数值仿真,分析了试验中的流动机理及试验结果差异产生的本质原因。提出了典型高超声速飞行器阻力预测准则。对飞行条件下的飞行器阻力进行预测,验证了预测准则的正确性。研究表明:壁温与来流静温比是造成不同风洞试验阻力差异的主要原因,对发动机内流道的压差阻力和摩擦阻力均有显著影响。在高超声速飞行器阻力预测时,要同时模拟马赫数、雷诺数、壁温与来流静温比3个相似参数。
In order to study the effect of wall temperature on the hovercraft hypersonic vehicle, experimental studies were conducted in conventional hypersonic wind tunnels and pulsed combustion wind tunnels. Numerical simulations were carried out to analyze the flow mechanism and the nature of the differences in test results. Proposed a typical hypersonic vehicle resistance prediction criteria. The prediction of the aircraft resistance under flight conditions proves the correctness of the prediction criterion. The results show that the ratio of wall temperature to flow and temperature is the main reason that causes the difference of resistance in different wind tunnel tests, and has significant influence on the pressure difference resistance and frictional resistance of the flow passage in the engine. In hypersonic vehicle resistance prediction, Mach number, Reynolds number, wall temperature and incoming flow static temperature ratio are simulated simultaneously with three similar parameters.