铱铑合金超声导波方法的固体火箭发动机燃烧室温度测试

IrRth40 Thermometry Combustion Chamber Temperature Measurement for Solid Rocket Motor

固体火箭发动机试车时温度参数是重要的测试物理量,国内外对于这种复杂环境的温度测试,除热电偶外尚无可靠的原位测试方法。为了研究固体火箭发动机试车时温度测试问题,用超声导波测温方法,设计出一套基于Ir Rth40(铱铑合金)超声导波测温系统,测试了该系统在室温~1600℃的运行情况。结果表明,超声测温系统可以在高温环境下稳定运行,并且室温~1600℃范围内校准曲线重复性良好。将获得的数据进行95置信度评估,绘制出95置信条件下的误差带。在温度大于1000℃时,灵敏度的变化幅度逐渐增大,达到0.0035μs/℃。常温常压下,传感器响应时间为1.2s。设计了传感器封装结构,完成了固体火箭发动机温度测试实验,测得温度-时间曲线,峰值温度为1492℃。

Solid Rocket Motor（SRM）temperature is an important physical parameter for which there is no reliable in-situ measurement within high temperature environments apart from the thermocouple. In this study,an ultrasonic temperature measurement system was designed with an Ir Rth40（iridium rhodium alloy）waveguide.Laboratory experiments showed that the device can obtain ultrasonic signals up to 1600℃ with a temperature fitting curve of room temperature to 1600℃. The thermometer also operated stably under high temperature and produced a repeatable calibration curve from room temperature to 1600℃. An error band was obtained via 95 confidence assessment. At temperatures above 1000℃,sensitivity gradually increased to a maximum of 0.0035μs/℃.The sensor response time was 1.2 s under normal temperature and pressure. A corresponding application structure was established for SRM before subjecting the sensor to a temperature test experiment. The temperature-time curve was obtained accordingly and peak temperature was measured at 1492℃.