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对比分析了两种气流状态参数和两种加热情况下典型前缘部件表面热流密度的相似性,论证了利用亚声速高温燃气流加热方式进行近地空间高超声速飞行工况气动热模拟试验的可行性.针对高超声速飞行器典型钝头锥结构提出“小喷口低速高温燃气流+石英灯”组合热试验方案.通过采用新型高效双腔蒸发管型燃气发生器、新型带保温夹层和耐高温陶瓷内衬的水冷不锈钢高温管道结构,同时引入电加热器预热及燃烧室两路供油方案,使所建低速高温燃气流热试验设备产生燃气流温度达到2 100K,250mm喷口处平均径向温度分布梯度约3K/mm,具有线性温度控制功能且稳态控制温差约46K,满足24km、马赫数为6典型高超声速飞行器工况驻点区域高温/大热流密度气动热试验要求.
The similarities between the two airflow state parameters and the surface heat flux density of the typical leading edge components are compared and analyzed. It is demonstrated that the subsonic hypersonic gas stream heating method is feasible for the aerodynamic thermal simulation of near-space hypersonic flight conditions Aiming at the typical blunt nose cone structure of hypersonic vehicle, this paper proposes a combined heat test scheme of “small nozzle high temperature gas stream + quartz lamp” .Through the adoption of a new type of high efficient double chamber evaporative tube gas generator, a new type of insulation sandwich with high temperature Ceramic lined water-cooled stainless steel high-temperature pipe structure, while the introduction of electric heater preheating and two combustion chamber fuel supply scheme, so that the built low-speed high-temperature gas flow thermal test equipment to produce gas flow temperature reaches 2 100K, 250mm nozzle mean diameter Temperature gradient of about 3K / mm, with linear temperature control and steady-state temperature control of about 46K, to meet the 24km, Mach number 6 typical hypersonic stagnation zone high temperature / high heat flux density aerodynamic test requirements.