Analysis of data correlation between combustion heated impulse facility and hypersonic wind tunnel
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摘要: 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。Abstract: 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.
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图 3 高超标模不同总温试验CA~α曲线
Figure 3. CA~α graph of different T0 tests for the typical hypersonic model
图 4 高超标模不同总温试验CN~α曲线
Figure 4. CN~α graph of different T0 tests for the typical hypersonic model
图 5 不通气标模轴向力系数计算和试验结果
Figure 5. Numerical and experimental axis force coefficients of the test model
图 6 不通气标模法向力系数计算和试验结果
Figure 6. Numerical and experimental normal force coefficients of the test model
图 7 CD~α试验与计算结果对比(脉冲风洞，总温1500K)
Figure 7. CD~α comparison of numerical and experimental results from Φ600mm combustion heated impulse facility
图 8 CL~α试验与计算结果对比(脉冲风洞，总温1500K)
Figure 8. CL~α comparison of numerical and experimental results from Φ600mm combustion heated impulse facility
图 9 CD~α试验与计算结果对比(常规高超声速风洞)
Figure 9. CD~α comparison of numerical and experimental results from Φ1m hypersonic wind tunnel
图 10 CL~α试验与计算结果对比(常规高超声速风洞)
Figure 10. CL~α comparison of numerical and experimental results from Φ1m hypersonic wind tunnel
图 11 CD~α计算与试验结果对比(脉冲燃烧风洞和常规高超声速风洞)
Figure 11. CD~α comparison of numerical and experimental results from combustion heated impulse facility and hypersonic wind tunnel
表 1 不同总温高超标模试验参数
Table 1. Flow parameters of different T0 tests
Ma p0/MPa T0/K p/Pa T/K Re 6.0 4.5 1200 2368 162 7.08×106 6.0 5.1 1500 2398 210 4.98×106 6.0 5.8 1800 2374 263 3.61×106 
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表 2 试验参数
Table 2. Test parameters
Facility Ma T0/K p0/MPa Φ2.4m CHIF 6 1575 4.76 Φ600mm CHIF 6 1500 4.61
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表 3 Φ1m常规高超声速风洞与Φ600mm脉冲燃烧风洞流场参数
Table 3. Test parameters of Φ1m hypersonic wind tunnel and Φ600mm combustion heated impulse facility
Facility Ma p0/MPa T0/K γ Re Tratio Φ1m HWT 6.0 3.00 500 1.40 2.62×107 5.17 Φ600mm CHIF 6.0 4.65 1560 1.36 4.10×106 1.49
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