脉冲燃烧风洞与常规高超声速风洞数据相关性研究

Analysis of data correlation between combustion heated impulse facility and hypersonic wind tunnel

不同风洞因模拟来流参数不同,对高超声速飞行器气动力试验结果影响很大。总结了脉冲燃烧风洞和常规高超声速风洞不通气标模的试验和计算结果,分析了水凝结、雷诺数、壁温比对模型气动性能的影响规律。脉冲燃烧风洞获得的气动性能变化规律与常规高超声速风洞一致,脉冲燃烧风洞获得的阻力系数比常规高超声速风洞阻力系数大15%左右,其中雷诺数影响较小,在5%以内,壁温比影响较大,在10%以上。结合数值计算对造成差异的原因进行分析,认为壁面传热对边界层速度型的影响是主要因素。

The simulation parameters of different hypersonic test facilities have large influence to the test results.A typical lift body aerodynamic test was conducted in the combustion heated impulse facility and hypersonic wind tunnel.The influence of Reynolds number and wall temperature ratio was analyzed.The aerodynamic changes of the combustion heated impulse facility test were in consistence with those of the hypersonic wind tunnel test,but its drag coefficients were about 15%larger.Numerical analysis indicates that the differences caused by Reynolds number were about 5%,and those caused by the wall temperature ratio were about 10%,and therefore the change of velocity boundary layer is the main factor.