Infrared thermography experiments of hypersonic boundary-layer transition on a 7° half-angle sharp cone
-
摘要: 为了推动高超声速边界层转捩研究的深入开展，给边界层转捩机理研究、物理模型验证、转捩数据库构建和转捩天地相关性的建立等提供基础风洞实验数据，在中国空气动力研究与发展中心的Φ1 m高超声速风洞开展了边界层转捩规律红外热图实验。针对半锥角7°尖锥模型，研究了不同单位雷诺数、迎角和马赫数对尖锥边界层转捩位置的影响规律。实验单位雷诺数（0.49~2.45）×107/m，迎角范围-10°~10°，马赫数5~7，模型头部半径0.05 mm。通过红外热图技术测量模型表面温度分布，获得了较为详细的转捩位置和转捩参数影响规律。实验结果表明：在马赫数5~7范围内，马赫数增大，尖锥转捩位置提前，分析认为是高马赫数时的雷诺数较大、自由流噪声水平较高引起；随着单位雷诺数的增大，边界层转捩位置前移，转捩雷诺数保持不变，约为3.0×106；小迎角时，随着迎角的增大，迎风面边界层转捩推迟，背风面边界层转捩前移，在10°大迎角时，迎风区中心线转捩前移，出现迎角"转捩逆转"现象，背风区出现了流动分离导致的低热流条带。Abstract: In order to promote the in-depth research on the hypersonic boundary layer transition and provide basic wind tunnel experimental data for the study of the boundary layer transition mechanism, the physical model validation, and the transition database construction, infrared thermography experiments of boundary layer transition are carried out in the Φ1 m hypersonic wind tunnel at CARDC. The effects of different unit Reynolds numbers, angles of attack and Mach numbers on the transition positions are studied on a 7° half-angle sharp cone. Test unit Reynolds numbers range from 0.49×107/m to 2.45×107/m. Test angles of attack range from -10° to 10°. Test Mach numbers range from 5 to 7. The head radius of the test model is 0.05mm. The quantitative infrared thermography technique is employed to obtain the temperature distribution photos of the model surface. By this way, the accurate transition positions and the effects of transition factors are obtained. Test results of the global temperature distribution show that an earlier transition occurs with the increase of Mach number. This is due to the larger Reynolds number and stronger flow field noise brought by the higher Mach number. As the unit Reynolds number increases, the transition position of the boundary layer moves forward and the transition Reynolds number remains constant about 3.0×106. When the angle of attack is small, a delayed transition occurs on the windward side and an earlier transition occurs on the leeward side with the increasing angle of attack. When the angle of attack is 10°, an earlier transition occurs at the center line of the windward side and reversed transition with angle of attack takes place, accompanied with a low heat flow strip induced by the flow separation on the leeward side.
-
Keywords:
- hypersonic /
- sharp cone /
- boundary layer transition /
- infrared thermography /
- Mach number /
- Reynolds number /
- angle of attack
-
-
![]()
图 3 模型表面温升分布(Re∞=1.0×107/m)
Fig. 3 Distribution of surface temperature rise(Re∞=1.0×107/m)
![]()
图 9 模型背风面温升分布(Rn=0.05，α=0° ~10°)
Fig. 9 Temperature rise on leeward side (Rn=0.05, α=0° ~10°)
![]()
图 11 转捩雷诺数随迎角变化关系(Rn=0.05，α=0° ~10°)
Fig. 11 Relationship between transition Reynolds number and angles of attack(Rn=0.05, α=0°~10°)
表 1 实验状态
Table 1 Test conditions
编号 Ma∞ Re∞ /m-1 α /(°) 研究内容 1 6 (0.49~2.45)×107 0 雷诺数对转捩的影响 2 6 1.0×107 -10~10 迎角对转捩的影响 3 5~7 1.0×107 0 马赫数对转捩的影响 
下载: 导出CSV
表 2 流场参数
Table 2 Parameters of flow field
Ma∞ p0/MPa T0/K p∞/Pa T∞/K Re∞/m-1 p'/Pa p'/p 5 0.50 384 945.02 64.00 0.96×107 49.9 2.8% 6 0.51 461 323.00 56.22 0.49×107 36.6 6.1% 6 0.78 472 494.00 57.56 0.72×107 55.5 5.0% 6 1.10 474 696.70 57.81 1.00×107 57.1 3.9% 6 2.82 488 1786.00 59.51 2.45×107 114.2 3.3% 7 2.49 594 601.47 55.00 1.09×107 95.2 6.5%
下载: 导出CSV
表 3 尖锥表面转捩位置测量结果(不同来流马赫数)
Table 3 Test results of transition position
Ma∞ Re∞
/m-1转捩位置
/mm转捩区长度
/mm转捩雷诺数
/m-1xT1 xT2 LT RexT 5 0.96×107 400 540 140 4.03×106 6 1.0×107 300 500 200 3.02×106 7 1.09×107 260 450 190 2.86×106
下载: 导出CSV
表 4 模型表面转捩位置测量结果(变雷诺数)
Table 4 Test results of transition position
Ma∞ Re∞
/m-1转捩位置
/mm转捩区长度
/mm转捩雷诺数
/m-1xT1 xT2 LT RexT 6 0.49×107 600 >800 >200 2.96×106 6 0.72×107 420 620 200 3.05×106 6 1.0×107 300 500 200 3.02×106 6 2.45×107 < 165 270 < 4.07×106
下载: 导出CSV
表 5 模型表面转捩位置测量结果(变迎角)
Table 5 Test results of transition position
Ma∞ Re∞
/m-1α/(°) 转捩位置
/mm转捩区长度
/mm转捩雷诺数
/m-1xT1 xT2 LT RexT1 6 0.97×107 10 < 165 < 165 < 1.61×106 6 0.92×107 6 < 165 < 165 < 1.53×106 6 0.94×107 2 < 165 < 165 < 1.56×106 6 1.00×107 0 300 500 200 3.02×106 6 0.97×107 -2 520 700 180 5.08×106 6 0.95×107 -4 600 770 170 5.74×106 6 0.97×107 -6 620 >800 >180 6.06×106 6 0.96×107 -8 未转捩 未转捩 >7.74×106 6 0.99×107 -10 660 >800 >140 6.58×106
下载: 导出CSV
-
[1] 陈坚强, 涂国华, 张毅锋, 等.高超声速边界层转捩研究现状与发展趋势[J].空气动力学学报, 2017, 35(3): 311-327. DOI: 10.7638/kqdlxxb-2017.0030 CHEN J Q, TU G H, ZHANG Y F, et al. Hypersonic boundary layer transition: what we know, where shall we go[J]. Acta Aerodynamica Sinica, 2017, 35(3): 311-327. DOI: 10.7638/kqdlxxb-2017.0030
[2] 柳森, 王宗浩, 谢爱民, 等.高超声速锥柱裙模型边界层转捩的弹道靶实验[J].实验流体力学, 2013, 27(6): 26-31. DOI: 10.3969/j.issn.1672-9897.2013.06.005 LIU S, WANG Z H, XIE A M, et al. Ballistic range experiments of hypersonic boundary layer transition on a cone-cylinder-flare configuration[J]. Journal of Experiments in Fluid Mechanics, 2013, 27(6): 26-31. DOI: 10.3969/j.issn.1672-9897.2013.06.005
[3] 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-2018.0017 LIU 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
[4] 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.1000-1328.2015.11.014 CHANG 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
[5] MUIR J F, TRUJILLO A 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[R]. AIAA-1972-216, 1972.
[6] STETSON K F, GEORGE H R. Shock tunnel investigation of boundary layer transition at Ma=5.5[J]. AIAA Journal, 1967, 5(5): 899-906.
[7] GROSSIR G, PINNA F, BONUCCI G, et al. Hypersonic boundary layer transition on a 7 degree half-angle cone at Mach 10[R]. AIAA-2014-2779, 2014.
[8] 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.
[9] WILLEMS S, GVLHAN 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.
[10] 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.
[11] HORVATH T J, BERRY S A. HOLLIS B R, et al. Boundary layer transition on slender cones in conventional and low disturbance Mach 6 wind tunnels[R]. AIAA-2002-2743, 2012.
[12] BI Z X, ZHU N J, SHEN Q, et al. Measurements on transition of hypersonic boundary layer over circular cone[R]. AIAA-2007-6728, 2007.
[13] ZHANG C H, TANG Q, LEE C B. Hypersonic boundary-layer transition on a flared cone[J]. Acta Mechanica Sinica, 2013, 29(1): 48-54. DOI: 10.1007/s10409-013-0009-2
[14] SOFTLEY E J. Boundary layer transition on hypersonic blunt, slender cones[R]. AIAA-1969-705, 1969.
[15] KIMMEL R L. The effect of pressure gradients on transition zone length in hypersonic boundary layers[J]. Journal of Fluids Engineering, 1997, 119(1): 36-41. http://www.wanfangdata.com.cn/details/detail.do?_type=perio&id=5c4f282494cda00935103a30cf644bec
[16] ANDERSON J D JR. Hypersonic and high temperature gas dynamics[M]. New-York: McGraw-Hill Book Company, 1989.
[17] OWEN F K, HORSTMAN C C, STAINBACK P C, et al. Comparison of wind tunnel transition and freestream disturbances measurements[J]. AIAA Journal, 1975, 13(3): 266-269. DOI: 10.2514/3.49691
[18] JULIANO T J, BORG M P, SCHNEIDER S P. Quiet tunnel measurements of HIFiRE-5 boundary-layer transition[J]. AIAA Journal, 2015, 53(4): 832-846. DOI: 10.2514/1.J053189
计量
- 文章访问数: 378
- HTML全文浏览量: 177
- PDF下载量: 56
