凹腔尺寸对迎风凹腔与逆向喷流组合热防护系统性能的影响

陆海波; 刘伟强

凹腔尺寸对迎风凹腔与逆向喷流组合热防护系统性能的影响

陆海波^1,2,
刘伟强¹

1.
国防科学技术大学航天与材料工程学院,长沙 410073
2.
中国人民解放军南京炮兵学院,南京 211100

基金项目: 国家自然科学基金(90916018); 高等学校博士学科点专项科研基金(200899980006)

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出版历程
- 收稿日期: 2011-11-23
- 刊出日期: 2012-12-28

Effect of cavity physical dimension on forward-facing cavity and opposing jet thermal protection system cooling efficiency

LU Hai-bo^1,2,
LIU Wei-qiang¹

1.
College of Aerospace and Material Engineering, National University of Defense Technology, Changsha 410073, China
2.
Nanjing Artillery Academy, The Chinese People's Liberation Army, Nanjing 211100, China

摘要

摘要: 针对高超声速飞行器热防护系统(TPS)的设计,对迎风凹腔与逆向喷流组合热防护系统展开研究.在数值方法实验验证的基础上,通过求解Navier-Stokes方程得到了带组合热防护系统的鼻锥的流场结构以及壁面热流分布.验证了组合热防护系统的有效性.在逆向喷流条件不变的情况下,进一步研究了凹腔的尺寸变化对其防热能力的影响.研究发现:凹腔的直径越小,深度越深,气动加热值越低.自由来流与逆向喷流形成的回流区在减少鼻锥的气动加热上起到关键的作用.相对于凹腔深度的变化,鼻锥壁面的气动加热更敏感于凹腔直径的变化.
- 热防护 /
- 高超声速飞行器 /
- 逆向喷流 /
- 迎风凹腔 /
- 数值模拟 /
- 气动加热
Abstract: Design of the thermal protection system (TPS) with the forward-facing cavity and opposing jet combined configuration was investigated numerically for the hypersonic vehicle.The numerical method was validated with the related experiments.The flow field parameters and surface heat flux distribution were obtained by solving the Navier-Stokes (N-S) equations.The validity of the combined TPS was testified and the effect of the cavity physical dimension on cooling efficiency of the combined TPS was discussed.The results show that the TPS with smaller diameter cavity has smaller aerodynamic heating and the TPS with larger length cavity has higher heat flux reduction.The recirculation region plays a pivotal role for the reduction of heat flux.The aerodynamic heating is more sensitive to the changing of the cavity diameter than the cavity length.
- thermal protection /
- hypersonic vehicle /
- opposing jet /
- forward-facing cavity /
- numerical simulation /
- aerodynamic heating

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参考文献(21)

[1]	Aiichiro T,Hiroyuki Y.Advanced thermal protection systems for reusable launch vehicles[R].AIAA-2001-1909, 2001.
[2]	王保国,李翔,黄伟光.激波后高温高速流场中的传热特性研究[J].航空动力学报,2010,25(5):963-980.WANG Baoguo,LI Xiang,HUANG Weiguang.Analytical study of heat transfer of high-temperature and high-velocity flow field behind shock waves[J].Journal of Aerospace Power,2010,25(5):963-980.(in Chinese)
[3]	侯玉柱,郑京良,董威.高超声速飞行器瞬态热试验[J].航空动力学报,2010,25(2):343-347.HOU Yuzhu,ZHENG Jingliang,DONG Wei.Transient test of aerodynamic heating for hypersonic vehicle[J].Journal of Aerospace Power,2010,25(2):343-347.(in Chinese)
[4]	Glass D E.Heat-pipe-cooled leading edges for hypersonic vehicles[R].Santa Barbara U.S.:Workshop on Materials and Struetres for Hypersonic Flight,2006.
[5]	叶宏,耿雪.热光伏技术在飞行器再入过程中的应用[J].中国科学:技术科学,2010,41(1):102-108.YE Hong,GENG Xue.The feasibility analysis of the application of TPV system in reentry[J].Scientia Sinica Techologica,2010,41(1):102-108.(in Chinese)
[6]	Warren C H E.An experimental investigation of the effect of ejecting a coolant gas at the nose of a bluff body[J].Journal of Fluid Mechanics,1960,8(3):400-417.
[7]	Aso S,Hayashi K,Mizoguchi M.A study on aerodynamic heating reduction due to opposing jet in hypersonic flow[R].AIAA-2002-0646,2002.
[8]	Hayashi K,Aso S.Effect of pressure ratio on aerodynamic heating reduction due to opposing jet[R].AIAA-2003-4041,2003.
[9]	Hayashi K, Aso S,Tani Y.Numerical study of thermal protection system by opposing jet [R].AIAA-2005-188,2005.
[10]	Tamada I,Aso S,Tani Y.Reducing aerodynamic heating by the opposing jet in supersonic and hypersonic flows[R].AIAA-2010-991,2010.
[11]	Hartmann J,Troll B.On a new method for the generation of sound waves[J].Physics Review,1922,20(6):719-727.
[12]	Engblom W A,Goldstein D B.Fluid dynamics of hypersonic forward-facing cavity flow[J].Journal of Spacecraft and Rockets,1997,34(4):437-444.
[13]	Burbank P B,Stallings R L.Heat-transfer and pressure measurements on a flat nose cylinder at a Mach number range of 2.49 to 4.44[R].NASA TM X-221,1959.
[14]	Yuceil B,Dolling D S,Wilson D.A preliminary investigation of the Helmholtz resonator concept for heat flux reduction[R].AIAA-1993-2742,1993.
[15]	Silton S I,Goldstein D B.Modeling of nose tip ablation onset in unsteady hypersonic flow[R].AIAA-2000-0204,2000.
[16]	Silton S I,Goldstein D B.Use of an axial nose-tip cavity for delaying ablation onset in hypersonic flow[J].Journal of Fluid Mechanics,2005,528(7):297-321.
[17]	Saravanan S,Nagashetty K,Jagadeesh G,et al.Experimental investigation of heat transfer reduction using forward facing cavity for missile shaped bodies flying at hypersonic speed[C]//26th International Symposium on Shock Waves Part Ⅷ.Gttingen Gemony:Springer,2007:613-617.
[18]	Saravanan S,Jagadeesh G,Reddy K P J.Investigation of missile-shaped body with forward-facing cavity at Mach 8[J].Journal of Spacecraft and Rockets,2009,46(3):577-591.
[19]	陶文铨.数值传热学[M].2版.西安:西安交通大学出版社,2001.
[20]	Baker A J.Numerical grid generation techniques[R].NASA CP-2166,1980.
[21]	Thomas P D,Middelcoff J F.Direct control of the grid point distribution in meshes generated by elliptic equations[R].AIAA-83-0037,1983.