摘要
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。
To study the effect of wall temperatures on drag of hypersonic vehicle,experiments are carried in a pulse combustion wind tunnel and a general hypersonic wind tunnel using a typical hypersonic vehicle model.Combined with simulation results,the flow mechanism is analyzed and the differences of the test results in each wind tunnel are investigated.The drag prediction principle of typical hypersonic vehicle is proposed.The validity is verified by the drag prediction results at flight condition.It is shown that the drag differences of the test results at wind tunnels are due to the ratio of wall temperature and inflow static temperature,which has an important effect on the pressure drag and the friction drag of the engine inner channel.When the drag of hypersonic vehicle is predicted,there are three similarity parameters which need to be simulated,i.e.,the Mach number,Reynolds number,as well as the ratio of wall temperature to inflow static temperature.
出处
《航空学报》
EI
CAS
CSCD
北大核心
2016年第4期1083-1091,共9页
Acta Aeronautica et Astronautica Sinica
基金
国家自然科学基金(91216102)~~
关键词
高超声速
壁温
气动阻力
数值模拟
风洞试验
hypersonic
wall temperature
aerodynamic drag
numerical simulation
wind tunnel test
