利用200 MW高温燃气流风洞,开展1:1全尺寸舱段模型热结构/热匹配试验.围绕风洞试验状态调节、介质影响热流设计评估、多台阶长时间热流校测等技术问题,开展高温燃气流风洞舱段类模型试验设计方法研究.引入新的燃气流风洞试验状态调节方法和考虑燃气介质差异的等效冷壁热流修正方法,并通过长时间变工况表面热流测试方法的建立,实现了多工况同步流场校测.试验设计方法有效应用于全尺寸舱段模型热结构/热匹配试验研究中.经试验验证,所建立的长时间大功率燃气流风洞试验设计方法有效可行,实现了全尺寸舱段飞行热环境的有效模拟.

为保证200 MW燃气流风洞高压氧气系统安全运行,从初始能量出发,对高压氧气系统充气、供气、排气时管道内的激波管流动、绝热压缩等过程进行安全分析,并提出针对性安全措施.结果表明:对于充气管道内存在的激波管流动,当驱动气体压力为20 MPa、被驱动气体压力为0.1 MPa时,激波反射后末端气体温度远远高于200℃,通过减小阀门开启速度,对阀前管道进行充气以减小上下游压差,可避免因绝热压缩产生的高温;供气管道充填时,管道内最高温度为73℃,通过控制充填速度,可进一步降低管道内氧气温度;通过高压排气、低压排气2种模式,可满足国标中对氧气流速的要求.研究结果可为氧气管道远程安全操作提供参考.

For the long time steady measurement of heat flux of MW magnitude, a calorimetric heat-flux sensor of new structure is designed based on the boiling heat transfer method. A U-shaped water cooling channel is designed, with expected measuremente range from 5 to 15 MW/m2. The sensor is tested in the high temperature supersonic gas flow field, to find its real range of 3 to 25 MW/m2. It is shown that the application of the boiling heat transfer method makes the design much easier due to the better heat transfer efficiency at a lower speed and a smaller pressure loss of the subcooled water, leading to its extended range for measurment.

This paper introduces a new thermal-dynamic coupled test method for spherical cone radome in high Temperature and supersonic jet flow field produced by rocket motor. Aerodynamic force of jet flow is insufficient to simulate transverse aerodynamic force of the radome at large angles of attack and maneuver flight trajectory, so the thermal-dynamic coupled test facility is designed. Test results show that thermal-dynamic coupled test is feasible in high temperature and supersonic jet flow field, the test method can be used for following test.