Pressure fluctuation experiments of hypersonic boundary-layer on a 7-degree half-angle sharp cone
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摘要: 针对半锥角7°尖锥模型，在常规高超声速风洞中开展了边界层脉动压力测量实验，进行了线性稳定性分析，研究了单位雷诺数和马赫数对尖锥边界层转捩位置和边界层稳定性的影响规律。模型长度800 mm，头部半径0.05 mm，实验单位雷诺数0.49 × 107～2.45 × 107 m–1，马赫数5～8，迎角0°。通过红外热图技术和高频脉动压力测量技术获得了模型表面边界层转捩位置和边界层内扰动波能谱分布，利用线性稳定性理论分析了最不稳定波频率和增长率。实验结果表明：在转捩区间可以测量到明显具有不稳定波频谱特征的脉动压力信号，其频率与稳定性理论分析的二模态不稳定波接近，幅值变化趋势也与之类似；随着雷诺数增大，不稳定波出现位置提前，频率增大，转捩位置提前；边界层中不稳定波包含第一和第二模态，马赫数5时，转捩由第一模态主导，马赫数高于6时，由第二模态主导。Abstract: In a conventional hypersonic wind-tunnel, pressure fluctuations of the boundary layer on a 7-degree half-angle sharp cone are measured by surface sensors and are analyzed by the linear stability theory. The influences of unit Reynolds numbers and Mach number on the stability and transition position of the boundary layer are studied. The length of the test model is 800 mm and the radius of the head is 0.05 mm. Test unit Reynolds numbers range from 0.49 × 10 7 m–1 to 2.45 × 107 m–1. Test Mach numbers range from 5 to 8. The angle of attack is 0°. The transition position and the energy spectrum distribution of the disturbance wave in the boundary layer are obtained by the quantitative infrared thermography and high frequency surface pressure fluctuation measurement techniques. The frequency and growth rate of the most unstable wave are analyzed by using the linear stability theory. The experimental results show that the fluctuating pressure signal with obvious characteristics of the unstable wave spectrum can be measured in the transition region. The frequency of the pressure fluctuation is close to that of the second mode instability analyzed by the linear stability theory, and the amplitude variation trend is also similar to that of the theoretical analysis. With the increase of the unit Reynolds number, the instability appears earlier, the dominant frequency is increased, and the transition onset moves forward. The unstable wave in the boundary layer contains the first and second modes. When the free-stream Mach number is equal to 5, the transition is caused by the first mode, and when the Mach number is above 6, the transition is attributed to the second mode.
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图 3 模型表面温升分布和脉动压力功率谱 （Re∞ = 1.00 × 107 m–1）
Figure 3. The temperature rise distribution and fluctuating pressure power spectrum of the model surface（Re∞ = 1.00 × 107 m–1）
图 4 模型表面温升沿子午线变化曲线
Figure 4. Surface temperature rise of centre lines on model surface
图 6 脉动压力功率谱和线性稳定性分析 （Re∞ = 0.49 × 107 m–1）
Figure 6. Fluctuating pressure power spectra and linear stability analysis （Re∞ = 0.49 × 107 m–1）
图 7 脉动压力功率谱和线性稳定性分析 （Re∞ = 0.72 × 107 m–1）
Figure 7. Fluctuating pressure power spectra and linear stability analysis （Re∞ = 0.72 × 107 m–1）
图 8 脉动压力功率谱和线性稳定性分析（Re∞ = 2.45 × 107 m–1）
Figure 8. Fluctuating pressure power spectra and linear stability analysis （Re∞ = 2.45 × 107 m–1）
图 9 模型表面温升沿子午线变化曲线
Figure 9. Surface temperature rise of centre lines on model surface
表 1 PCB传感器安装位置
Table 1. PCB installation locations
PCB
编号1 2 3 4 5 6 7 8 l/mm 125 205 285 365 445 525 605 685 
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[1] 罗纪生. 高超声速边界层的转捩及预测[J]. 航空学报, 2015, 36(1): 357–372.LUO J S. Transition and prediction for hypersonic boundary layers[J]. Acta Aeronautica et Astronautica Sinica, 2015, 36(1): 357–372. [2] SMITH F T. On the first-mode instability in subsonic, supersonic or hypersonic boundary layers[J]. Journal of Fluid Mechanics, 1989, 198: 127–153. doi: 10.1017/s0022112089000078 [3] MACK L M. Linear stability theory and the problem of supersonic boundary-layer transition[J]. AIAA Journal, 1975, 13(3): 278–289. doi: 10.2514/3.49693 [4] MACK L M. Boundary-layer linear stability theory[R]. AGARD Report 709, 1984. [5] SARIC W S, Reed H L. Crossflow instabilities - theory & technology[R]. AIAA 2003-771, 2003. doi: 10.2514/6.2003-771 [6] SARIC W S. Görtler vortices[J]. Annual Review of Fluid Mechanics, 1994, 26(1): 379–409. doi: 10.1146/annurev.fl.26.010194.002115 [7] 刘向宏, 赖光伟, 吴杰. 高超声速边界层转捩实验综述[J]. 空气动力学学报, 2018, 36(2): 196–212. doi: 10.7638/kqdlxxb-2018.0017LIU X H, LAI G W, WU J. Boundary-layer transition experiments in hypersonic flow[J]. Acta Aerodynamica Sinica, 2018, 36(2): 196–212. doi: 10.7638/kqdlxxb-2018.0017 [8] SCHMISSEUR J D. Hypersonics into the 21st century: a perspective on AFOSR-sponsored research in aerothermo-dynamics[J]. Progress in Aerospace Sciences, 2015, 72: 3–16. doi: 10.1016/j.paerosci.2014.09.009 [9] 沈清, 袁湘江, 王强, 等. 可压缩边界层与混合层失稳结构的研究进展及其工程应用[J]. 力学进展, 2012, 42(3): 252–261.SHEN Q, YUAN X J, WANG Q, et al. Review on the instability structure in compressible boundary layers and mixing layers and its application[J]. Advances in Mecha-nics, 2012, 42(3): 252–261. [10] MUIR J, TRUJILLO A. Experimental investigation of the effects of nose bluntness, free-stream unit Reynolds number, and angle of attack on cone boundary layer transition at a Mach number of 6[C]//Proc of the 10th Aerospace Sciences Meeting. 1972. doi: 10.2514/6.1972-216 [11] STETSON K F, RUSHTON G H. Shock tunnel investiga-tion of boundary-layer transition at M = 5.5[J]. AIAA Journal, 1967, 5(5): 899–906. doi: 10.2514/3.4098 [12] JULIANO T J, KIMMEL R L, WILLEMS S, et al. HIFiRE-1 boundary-layer transition: ground test results and stability analysis[R]. AIAA 2015-1736, 2015. doi: 10.2514/6.2015-1736 [13] WILLEMS S, GUELHAN A, JULIANO T J, et al. Laminar to turbulent transition on the HIFiRE-1 cone at Mach 7 and high angle of attack[R]. AIAA 2014-0428, 2014. doi: 10.2514/6.2014-0428 [14] JULIANO T J, KIMMEL R L, WILLEMS S, et al. HIFiRE-1 surface pressure fluctuations from high Reynolds, high angle ground test[R]. AIAA 2014-0429, 2014. doi: 10.2514/6.2014-0429 [15] STETSON K F, THOMPSON E R, DONALDSON J C, et al. Laminar boundary layer stability experiments on a cone at Mach 8, part 1: sharp cone[R]. AIAA-83-1761, 1983. [16] CASPER K M, BERESH S J, HENFLING J F, et al. Hypersonic wind-tunnel measurements of boundary-layer transition on a slender cone[J]. AIAA Journal, 2016, 54(4): 1250–1263. doi: 10.2514/1.j054033 [17] ZHANG C H, LEE C. Rayleigh-scattering visualization of the development of second-mode waves[J]. Journal of Visualization, 2017, 20(1): 7–12. doi: 10.1007/s12650-016-0384-4 [18] ZHU Y D, ZHANG C H, CHEN X, et al. Transition in hypersonic boundary layers: role of dilatational waves[J]. AIAA Journal, 2016, 54(10): 3039–3049. doi: 10.2514/1.j054702 [19] 常雨, 陈苏宇, 张扣立. 高超声速边界层转捩特性试验探究[J]. 宇航学报, 2015, 36(11): 1318–1323. doi: 10.3873/j.issn.1000-1328.2015.11.014CHANG Y, CHEN S Y, ZHANG K L. Experimental investigation of hypersonic boundary layer transition[J]. Journal of Astronautics, 2015, 36(11): 1318–1323. doi: 10.3873/j.issn.1000-1328.2015.11.014 [20] LIU X L, YI S H, XU X W, et al. Experimental study of second-mode wave on a flared cone at Mach 6[J]. Physics of Fluids, 2019, 31(7): 074108. doi: 10.1063/1.5103192 [21] 陈久芬, 凌岗, 张庆虎, 等. 7°尖锥高超声速边界层转捩红外测量实验[J]. 实验流体力学, 2020, 34(1): 60–66. doi: 10.11729/syltlx20180172CHEN J F, LING G, ZHANG Q H, et al. Infrared thermography experiments of hypersonic boundary-layer transition on a 7° half-angle sharp cone[J]. Journal of Experiments in Fluid Mechanics, 2020, 34(1): 60–66. doi: 10.11729/syltlx20180172 [22] CHEN X, ZHU Y D, LEE C. Interactions between second mode and low-frequency waves in a hypersonic boundary layer[J]. Journal of Fluid Mechanics, 2017, 820: 693–735. doi: 10.1017/jfm.2017.233 -
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