中文摘要：

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.

英文摘要：

Comparability of flow parameters and surface heal flux density on a typical leading structure by these two heating approach was analyzed,it was demonstrated to be feasible of utilizing lower-speed and high-temperature gas flow as heating method for hypersonic aerodynamic thermal test.A new combined heating scheme composed of a lower-speed and high-temperature gas flow and quartz lights was proposed for hypersonic aerodynamic test of blunt cone’s.The newly-built aerodynamic thermal tester was designed with an efficient double-chamber vaporizing combustor,a new water-cooled pipe structure with thermal insulation sandwich of keramic material bearable high temperature,electric pre-heaters and twin-fule-supply scheme for the combustor.It can generate a hot gas of 2 100 K.The average gradient of radial temperature distribution of 250 mm nozzle is about 3 K/mm.And the tester has auto-linear-control ability of temperature and the temperature difference is 46 K.In a word,it can meet the urgent requirement for thermal test of high temperature and high heat flux of hypersonic vehicle flying by 24 km height and mach number 6 speed.