前缘钝度对马赫数6平面压缩进气道流场的影响分析
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摘要
高超声速飞行要求前缘钝化,钝化前缘将引起流场变化并影响进气道性能,进气道几何参数需要进行修正设计。为了解钝化前缘影响的程度及条件,为修正设计提供依据,采用S-A模型,研究了马赫数6条件下,固定外罩前缘钝度(Rc=2mm),前体前缘钝化(RN=1~6mm)对一种3+1波系平面压缩进气道流场特征和综合性能的影响。结果表明,前体前缘钝化引起外压缩激波偏离唇口、肩部分离区向上游传播,造成分离激波与其他激波形成复杂的相交结构,进而形成严重畸变的唇口弓形激波,在较大钝度(RN≥3mm)时激波干扰生成的滑移层进入内流道核心区,唇口弓形激波的畸变和大范围的滑移层吞入,是造成总压恢复系数严重下降的主要原因;外压缩激波的外移以及分离激波的介入是造成流量捕获急剧下降的主要原因。由此提示,应当通过合理的钝化设计或流场修正设计,使前体前缘钝化的进气道避免外压缩激波偏离唇口太远,并通过流动控制方法严格控制肩部分离区范围,勿使分离激波与其他激波干扰产生的滑移层进入内流道核心区范围,更要避免分离激波与唇口弓形激波在内流道核心区域相交。
Leading-edges for hypersonic systems might be blunted,and forbody-inlet flowfield might be distortion and performance is deteriorated. The inlet geometry needs to be redesigned. To know critical factors and conditions before inlet redesign,S-A model is used to study forebody leading-edge bluntness effect on flowfield and properties in a 3+1 shock system planar inlet at Ma=6. Forebody leading-edge radius is from RN=1 to 6 mm with a cowl-lip radius of 2 mm. The results show that external compression shocks move away from cowl-lip and shoulder separated zone goes upstream with forebody leading-edge more blunted. The separation shock interacts with other shocks and cause complex flowfield. The cowl-lip shock is distorted seriously. When RN≥3 mm,shear layers from shock-shock interaction point are swallowed into the center zone of internal flow-path. The distorted cowl-lip shock and large scope of shear layers are critical factors to sharply decrease pressure recovery. The moving of external compression shocks and separation shock interaction lead to remarkable mass-flow decrease.These results suggest that reasonable bluntness and flowfield design is necessary to avoid external compression shocks moving far away cowl-lip. The separated zone size should be controlled strictly to avoid swallow of shear layers from the interaction point between separation shock and cowl-lip shock.
Leading-edges for hypersonic systems might be blunted,and forbody-inlet flowfield might be distortion and performance is deteriorated. The inlet geometry needs to be redesigned. To know critical factors and conditions before inlet redesign,S-A model is used to study forebody leading-edge bluntness effect on flowfield and properties in a 3+1 shock system planar inlet at Ma=6. Forebody leading-edge radius is from RN=1 to 6 mm with a cowl-lip radius of 2 mm. The results show that external compression shocks move away from cowl-lip and shoulder separated zone goes upstream with forebody leading-edge more blunted. The separation shock interacts with other shocks and cause complex flowfield. The cowl-lip shock is distorted seriously. When RN≥3 mm,shear layers from shock-shock interaction point are swallowed into the center zone of internal flow-path. The distorted cowl-lip shock and large scope of shear layers are critical factors to sharply decrease pressure recovery. The moving of external compression shocks and separation shock interaction lead to remarkable mass-flow decrease.These results suggest that reasonable bluntness and flowfield design is necessary to avoid external compression shocks moving far away cowl-lip. The separated zone size should be controlled strictly to avoid swallow of shear layers from the interaction point between separation shock and cowl-lip shock.
引文
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[13]周忠平.钝化对高超声速进气道性能的影响[D].南京:南京航空航天大学,2007.
[14]高文智,李祝飞,杨基明.一种鼻锥钝化高超声速轴对称进气道流动特性实验[J].航空学报,2015,36(1):302-310.
[15]张堃元.基于弯曲激波压缩系统的高超声速进气道反设计研究进展[J].航空学报,2015,36(1):274-288.
[16]潘瑾,张堃元,金志光.弯曲激波压缩型面的设计及数值分析[J].推进技术,2008,29(4):438-442.(PAN Jin,ZHANG Kun-yuan,JIN Zhi-guang.Design and Numerical Investigation of Curved Shock Compression Surface[J].Journal of Propulsion Technology,2008,29(4):438-442.)
[17]南向军,张堃元,金志光.矩形转圆形高超声速内收缩进气道数值及试验研究[J].航空学报,2011,32(6):988-996.
[18]贺旭照,周正,倪鸿礼.密切内锥乘波前体进气道一体化设计和性能分析[J].推进技术,2012,33(4):510-515.(HE Xu-zhao,ZHOU Zheng,NI Hong-li.Integrated Design Methods and Performance Analyses of Osculating Inward Turning Cone Waveride Forebody Inlet[J].Journal of Propulsion Technology,2012,33(4):510-515.)
[19]Wilson F N Santos.Bluntness Effects on Lift-to-Drag Ratio of Leading Edges for Hypersonic Waverider Configurations[R].AIAA 2012-5802.
[20]李程鸿,谭慧俊,孙姝.流体式高超声速可调进气道流动机理及工作特性分析[J].宇航学报,2011,32(12):2613-2621.
[21]张悦,谭慧俊,张启帆,等.一种进气道内激波/边界层干扰控制的新方法及其流动机理[J].宇航学报,2012,33(12):265-274.
[22]Holger Babinsky,John K Harvey.Shock Wave-Boundary-Layer Interactions[M].New York:Cambridge University Press,2011:11-16.
[23]John D Anderson.Hypersonic and High-Temperature Gas Dynamics[M].Washington DC:American Institute of Aeronautics and Astronautics,2006.
[24]John J Bertin.Hypersonic Aerothermodynamics[M].Washington DC:American Institute of Aeronautics and Astronautics,1994.
[2]约翰.冲压发动机技术(下)[M].李存杰,付巽权,译.北京:国防工业出版社,1987.
[3]Corin Segal.The Scramjet Engine Progress and Characteristics[M].New York:Cambridge University Press,2009.
[4]Sheldon B Cousin.Leading-Edge Bluntness Effects and Their Importance in Hypersonic Inlet Design[R].AIAA67-0451.
[5]张红军,沈清.高超声速进气道前缘钝度效应试验研究[J].推进技术,2013,34(10):1316-1320.(ZHANG Hong-jun,SHEN Qing.Experimental Studies of Leading-Edge Bluntness Effects on Hypersonic Inlet[J].Journal of Propulsion Technology,2013,34(10):1316-1320.)
[6]James M Cubbage.Effect of Nose Bluntness and Controlled Roughness on the Flow on Two Hypersonic Inlet Center Bodies with Cowling at Mach 5.98[R].NASA TN D-2900,1965.
[7]王卫星,郭荣伟.基于边界层转捩的高超声速进气道特性研究[J].航空学报,2012,33(10):1772-1780.
[8]王振锋.脉冲设备二维高超进气道压缩面激波-边界层干扰显示技术及应用研究[D].绵阳:中国空气动力研究与发展中心,2015.
[9]王振锋,白菡尘,桂业伟.前缘半径对两个尺度三级压缩楔流场结构影响研究[J].推进技术,2014,35(4):449-454.(WANG Zhen-feng,BAI Han-chen,GUI Yewei.Leading-Edge Radius and Scale Effect on Flowfield Structure of Three-Stage Compression Ramp[J].Journal of Propulsion Technology,2014,35(4):449-454.)
[10]王振锋,张扣立,江涛.三级压缩楔前缘半径对壁面静压及热流分布影响研究[J].推进技术,2014,35(8):1009-1015.(WANG Zhen-feng,ZHANG Kouli,JIANG Tao.Research on Leading-Edge Radius Effect on Wall Static Pressure and Heat Flux Distribution of Three-Stage Compression Ramp[J].Journal of Propulsion Technology,2014,35(8):1009-1015.)
[11]Benson J L,Maslowe S A.Bluntness and Boundary Layer Displacement Effects on Hypersonic Inlet Flow Fields[R].AIAA 65-0617.
[12]王晓栋,乐嘉陵.前缘对进气道性能影响的数值模拟[J].推进技术,2002,23(6):460-462.(WANG Xiao-dong,LE Jia-ling.Numerical Simulation of Effects of Leading-Edge on the Performance of Inlet[J].Journal of Propulsion Technology,2002,23(6):460-462.)
[13]周忠平.钝化对高超声速进气道性能的影响[D].南京:南京航空航天大学,2007.
[14]高文智,李祝飞,杨基明.一种鼻锥钝化高超声速轴对称进气道流动特性实验[J].航空学报,2015,36(1):302-310.
[15]张堃元.基于弯曲激波压缩系统的高超声速进气道反设计研究进展[J].航空学报,2015,36(1):274-288.
[16]潘瑾,张堃元,金志光.弯曲激波压缩型面的设计及数值分析[J].推进技术,2008,29(4):438-442.(PAN Jin,ZHANG Kun-yuan,JIN Zhi-guang.Design and Numerical Investigation of Curved Shock Compression Surface[J].Journal of Propulsion Technology,2008,29(4):438-442.)
[17]南向军,张堃元,金志光.矩形转圆形高超声速内收缩进气道数值及试验研究[J].航空学报,2011,32(6):988-996.
[18]贺旭照,周正,倪鸿礼.密切内锥乘波前体进气道一体化设计和性能分析[J].推进技术,2012,33(4):510-515.(HE Xu-zhao,ZHOU Zheng,NI Hong-li.Integrated Design Methods and Performance Analyses of Osculating Inward Turning Cone Waveride Forebody Inlet[J].Journal of Propulsion Technology,2012,33(4):510-515.)
[19]Wilson F N Santos.Bluntness Effects on Lift-to-Drag Ratio of Leading Edges for Hypersonic Waverider Configurations[R].AIAA 2012-5802.
[20]李程鸿,谭慧俊,孙姝.流体式高超声速可调进气道流动机理及工作特性分析[J].宇航学报,2011,32(12):2613-2621.
[21]张悦,谭慧俊,张启帆,等.一种进气道内激波/边界层干扰控制的新方法及其流动机理[J].宇航学报,2012,33(12):265-274.
[22]Holger Babinsky,John K Harvey.Shock Wave-Boundary-Layer Interactions[M].New York:Cambridge University Press,2011:11-16.
[23]John D Anderson.Hypersonic and High-Temperature Gas Dynamics[M].Washington DC:American Institute of Aeronautics and Astronautics,2006.
[24]John J Bertin.Hypersonic Aerothermodynamics[M].Washington DC:American Institute of Aeronautics and Astronautics,1994.