7°尖锥高超声速边界层转捩红外测量实验  

Infrared thermography experiments of hypersonic boundary-layer transition on a 7° half-angle sharp cone

作　　者：陈久芬[1] 凌岗[1] 张庆虎 解福田[1] 许晓斌[1] 张毅锋[2] CHEN Jiufen;LING Gang;ZHANG Qinghu;XIE Futian;XU Xiaobin;ZHANG Yifeng(Hypervelocity Aerodynamics Institute of China Aerodynamics Research and Development Center,Mianyang Sichuan 621000,China;Computational Aerodynamics Institute of China Aerodynamics Research and Development Center,Mianyang Sichuan 621000,China)

机构地区：[1]中国空气动力研究与发展中心超高速空气动力研究所,四川绵阳621000 [2]中国空气动力研究与发展中心计算空气动力研究所,四川绵阳621000

出　　处：《实验流体力学》2020年第1期60-66,共7页Journal of Experiments in Fluid Mechanics

基　　金：国家重点研发计划(2016YFA0401200)。

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Φ1m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)×10~7/m,迎角范围-10°~10°,马赫数5~7,模型头部半径0.05mm。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3.0×10~6;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。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 0.49×10~7/m to 2.45×10~7/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 0.05 mm.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 3.0×10~6.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.

关 键 词：高超声速 尖锥 边界层转捩 红外热图 马赫数 雷诺数 迎角 

分 类 号：V211.7[航空宇航科学与技术—航空宇航推进理论与工程]