高超声速飞行器前体/冲压发动机一体化气动热实验研究

以超燃冲压发动机为动力的飞行器,由于飞行速度的增加,气动加热增强,而且在高马赫数范围内,冲压发动机燃烧室的滞止温度也是很高的.通过风洞实验,采用铂膜电阻温度计热流测量技术,开展了来流马赫数6.4和马赫数4.0两种状态下的热流分布规律研究,给出了前体、中支板及内通道的热流实验结果,研究了边界层流动状态、边界层抽吸、激波反射对热流分布的影响.实验结果表明,边界层流动状态对热流分布产生显著的影响,前体湍流热流值约为层流热流值的3.3倍;边界层抽吸会引起热流率增加;激波反射和激波加热对热流分布影响显著,马赫数越大激波加热越强.

Aeroheating experiment research on hypersonic integrated vehicle forebody/ramjet

A scramjet-powered aircraft,due to the increase in speed,aeroheating is very serious.Moreover,high flighting Mach number makes higher stagnation temperature within the ramjet combustor.In this text,heat flux measurements were conducted through wind tunnel experiments using platinum film resistance thermometer measurement technology,and experiment mach number is M=6.4 and M=4.0.Experiment results were presented about heat flux distribution on forebody,strut and flow path,which were used to study effects of boundary flow regime,suction and shockwave reflection on heat flux distribution.Results show that boundary layer flow regime has significant impact on heat flux distribution,and heat flux value of turbulence at ramp is about 3.3 times of laminar value.Boundary layer suction experiment reveals that suction increases rate of flux.Heat flux distribution and flow study show that shockwave heating and reflection markedly influence heat flux distribution,and high Mach number produces more serious shockwave heating.