受试验设备能力限制,地面风洞无法完全模拟高超声速飞行器临近空间热环境.文章采用在飞行器表面开孔安装长时耐高温热流传感器直接测量热流密度的方法,国内首次获得 M a12以上高超声速飞行器表面热流密度时变数据和边界层转捩特征.实测热流值与理论预示值规律相同,两者偏差小于20%.针对树脂基材料导热微分方程中虽考虑了热解吸热项,但未考虑导热系数随温度变化情况,采用在树脂基材料导热微分方程中加入物性参数随温度变化项的方法,计算了飞行器热防护结构内部分层温度和碳化层厚度,并与实测结果进行了比较,不考虑树脂热解特性和材料物性参数随温度变化,理论值高于实测值,最大偏差275～320℃;考虑热解特性和物性参数随温度变化情况,计算值与实测值最大偏差小于70℃.%Limited by the capacity of the test equipment ,the ground wind tunnel can′t fully simulate the near space thermal environment of the hypersonic vehicle . The longtime high temperature heat sensor was installed on the surface of the aircraft .For the first time in china ,the characteristics of heat flux density time-varying data and boundary layer transition in the hypersonic vehicle above Ma 12 was obtained . The measured heat flow value is the same as the theoretical prediction value ,and the deviation is less than 20% . Theoretical calculation of the internal temperature of resin based composites is higher than actual value , by the method of adding physical property parameters with temperature variation and the thermal conductivity in the heat conduction differential equation of resin matrix .And the internal stratified temperature and thickness of carbonization layer were calculated . Compared with the measured results ,the deviation between the calculated value and measured result is less than 70℃,but the calculated value is higher than measured result and the maximum deviation is 275～320℃,when the thermal conductivity and physical property parameters with temperature variation are not considered .
展开▼
