YuanChaokai1,2,*,, LiJinping1, ChenHong1, JiangZonglin1,2, YuHongru1 1State Key Laboratory of High-Temperature Gas Dynamics, Institute of Mechanics, Chinese Academy of Sciences, Beijing 100190, China2School of Engineering Sciences,University of Chinese Academy of Sciences, Beijing 100049, China 中图分类号:V211.7 文献标识码:10.6052/0459-1879-17-289
关键词:高超声速热防护;主动冷却;溢流冷却;液膜 Abstract Hypersonic overflow cooling is a new type of aircraft thermal protection method. The basic idea is that the overflow hole is arranged in the high heat flux area, and the coolant poured out in an overflow way.The liquid film spreads through the aircraft surface friction forming a thermal buffer layer to reduce the surface heat flux. Now, the overflow cooling technology is still in the exploratory stage, and a large number of experimental verification and mechanism research work need to do. In this paper, wind tunnel experiment platform for overflow cooling was build, adopting the heat flux measurement, liquid film thickness measurement and liquid film motion observation technology. The feasibility of applying overflow cooling to hypersonic thermal protection was verified, and the characteristics of liquid film flow under hypersonic flow field were preliminary analyzed. Reserch results show that: (1) In hypersonic flow field, the liquid film can be formed on the vehicle surface , and effectively isolate the external high temperature air to reduce the surface heat flux; (2) On wedge surface, the leading velocity of the liquid film gradually decelerate. Increase coolant flow rate, the liquid film thickness change is not obvious, but the leading velocity of liquid film will increase; (3) Surface waves exist in liquid film, and evolve in time and space direction, which leads to slight perturbation of liquid film thickness; (4) There is a lateral expansion phenomenon in the liquid film layer, that is, the width of the liquid film is greater than that of the overflow hole. The reason is that the liquid film layer don’t match the flow field boundary condition, and there is pressure gradient, forcing the coolant to flow to low pressure area, thus broadening the liquid film layer.
通过典型模型的溢流冷却风洞实验,对溢流冷却的热防护性能和液膜流动规律进行了初步研究,结论如下: (1) 高超声速条件下通过溢流能够在飞行器表面形成液膜并有效隔离外部高温气流,降低飞行器表面热流率,证明溢流冷却可以应用于高超声速飞行器热防护. (2) 楔面上的液膜前缘流动是一个逐渐减速的过程,增加冷却液流量液膜厚度变化不明显,但液膜前缘运动速度增加. (3) 液膜层存在表面波,在时间和空间方向发生演化,导致液膜厚度的微弱扰动. (4) 液膜层存在横向展宽现象,即液膜层宽度大于溢流缝宽度. 原因是液膜层与流场边界层条件不匹配,存在压力梯度,迫使冷却液向低压区流动,从而展宽液膜层,并且流量越高,横向展宽现象越明显. 高超声速条件下气液两相流可供参考的研究结果很少,而且高超声速风洞实验时间较短开展溢流冷却实验还存在诸多限制条件,还有待进一步深入研究液膜流动特性. The authors have declared that no competing interests exist.
LiuCheng, YeZhengyin, YeKun.The effect of transiton location on aerothermoelasticity of a hypersonic all-movable centrol surface . Chinese Journal of Theoretical and Applied Mechanics, 2017, 49(4): 802-810 (in Chinese) [本文引用: 1]
[2]
GlassDE.Ceramic matrix composite (CMC) thermal protection systems (TPS) and hot structures for hypersonic vehicles //15th AIAA Space Planes and Hypersonic Systems and Technologies Conference.2008, 28: 1-36 [本文引用: 1]
[3]
ShenL, WangJH, DongWJ, et al.An experimental investigation on transpiration cooling with phase change under supersonic condition . Applied Thermal Engineering, 2016, 105: 549-556 [本文引用: 1]
[4]
WangJH, ZhaoLJ, WangXC, et al.An experimental investigation on transpiration cooling of wedge shaped nose cone with liquid coolant . International Journal of Heat and Mass Transfer, 2014, 75: 442-449
[5]
HuangG, ZhuYH, LiaoZY, et al.Experimental investigation of transpiration cooling with phase change for sintered porous plates . International Journal of Heat and Mass Transfer, 2017, 114: 1201-1213 [本文引用: 1]
[6]
Barzegar GerdroodbaryM, ImaniM, GanjiDD.Investigation of film cooling on nose cone by a forward facing array of micro-jets in Hypersonic flow . International Communications in Heat and Mass Transfer, 2015, 64: 42-49 [本文引用: 1]
[7]
SriramR, JagadeeshG.Film cooling at hypersonic Mach numbers using forward facing array of micro-jets . International Journal of Heat and Mass Transfer, 2009, 52(15): 3654-3664
[8]
WangZG, SunXW, HuangW, et al.Experimental investigation on drag and heat flux reduction in supersonic/hypersonic flows: A survey . Acta Astronautica, 2016, 129: 95-110 [本文引用: 1]
HouYipeng, HouChi, WanXiaopeng, et al.Analysis of influence parameters to convection active cooling structure . Journal of Solid Rocket Technology, 2016, 39(1): 90-94 (in Chinese) [本文引用: 1]
SunJian, LiuWeiqiang.Research on convective cooling effect of leading edge platelet of airfoil . Acta Phys Sin, 2012, 61(12): 379-386 (in Chinese)
[11]
AnthonyF, DukesW, HelenbrookR.Data and results from a study of internal convective cooling systems for hypersonic aircraft, NASA CR-132432, 1974 [本文引用: 1]
[12]
AsoS, HayashiK, MizoguchiM.A study on aerodynamic heating reduction due to opposing jet in hypersonic flow . AIAA Paper, 2002, 646: 2002 [本文引用: 1]
[13]
HayashiK, AsoS, TaniY.Numerical study of thermal protection system by opposing jet//43rd Aiaa Aerospace Sciences Meeting and Exhibit, 2005: 2005-0188
HeKun, ChenJianqiang, DongWeizhong.Penetration mode and drag reduction research in hypersonic flows using a counter-flow jet . Chinese Journal of Theoretical and Applied Mechanics, 2006, 38(4): 438-445 (in Chinese)
[15]
FujitaM.Axisymmetric oscillations of an opposing jet from a hemispherical nose . AIAA Journal, 2015, 33(33): 1850-1856 [本文引用: 1]
GengYunfei, YanChao.Numerical Investigation of self-aligning spiked bodies at hypersonic speeds . Chinese Journal of Theoretical and Applied Mechanics, 2011, 43(3): 441-446 (in Chinese) [本文引用: 1]
[17]
AhmedMYM, QinN.Recent advances in the aerothermodynamics of spiked hypersonic vehicles . Progress in Aerospace Sciences, 2011, 47(6): 425-449
[18]
GauerM, PaullA.Numerical investigation of a spiked blunt nose cone at hypersonic speeds . Journal of Spacecraft & Rockets, 2015, 45(2008): 459-471
ZhangJiang, WuJunfei, NiWenbin, et al.Experimental investigation on flowfield around blunt body with forward-facing jet and spike . Chinese Journal of Theoretical and Applied Mechanics, 2016, 48(5): 1040-1048 (in Chinese) [本文引用: 1]
[20]
SiltonSI, GoldsteinDB.Use of an axial nose-tip cavity for delaying ablation onset in hypersonic flow . Journal of Fluid Mechanics, 2005, 528: 297-321 [本文引用: 1]
[21]
EngblomWA, GoldsteinDB.Nose-tip surface heat reduction mechanism . Journal of Thermophysics and Heat Transfer, 1996, 10(4): 598-606
[22]
SelvarajS, GopalanJ, ReddyKPJ.Investigation of missile-shaped body with forward-facing cavity at Mach 8 . Journal of Spacecraft and Rockets, 2009, 46(3): 577-591 [本文引用: 1]
ChenBing.Investigation on reducing the heat flux at the leading edge by overflow cooling. [PhD Thesis] . Beijing: University of Chinese Academy of Sciences, 2013 (in Chinese) [本文引用: 1]