激波风洞高焓流动及其驻点对流和辐射热流测量

HIGH ENTHALPY EQUILIBRIUM FLOW ANDMEASUREMENTS OF STAGNATION POINT CONVECTIVEAND RADIATIVE HEAT TRANSFER IN A SHOCK TUNNEL

在激波风洞中用氢氧燃烧驱动方法获得了总压１４ＭＰａ，总温高达７２００Ｋ的高超声速高焓平衡流，可以模拟再入飞行速度４至５ｋｍ／ｓ的真实气体效应。本文还介绍了高温气流中驻点对流和辐射传热测量技术及其测量结果。

The hydrogenoxygen combustion driver technique was used to produce strong shock waves and a high reservoir temperature range of 5800～7200K with a reservoir pressure of 14MPa.Hypersonic equilibrium flow in the conical nozzle was established with a duration of 4.5ms of fairly uniform flow at Mach number of 6.4 in the test section,in which real gas effects at reentry speed of 3～5km/s can be simulated. Convective and radiative heat transfer measurements were carried out at the stagnation point of a blunt model with copper slug calorimeters and thin film heat transfer radiation gages respectively.The results are compared with existing experimental data and theoretical prediction.