喷流落压比对高超飞行器尾喷管内外流干扰的实验  

Experiment of influence of the nozzle pressure ratio on the interaction between the external flow and nozzle flow of hypersonic aerocraft

作　　者：秦思[1] 贺旭照[1] 曾学军[1] 乐嘉陵[1] 

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学研究所高超声速冲压发动机技术重点实验室,四川绵阳621000

出　　处：《航空动力学报》2017年第10期2491-2497,共7页Journal of Aerospace Power

摘　　要：为了研究吸气式高超声速飞行器尾喷流对飞行器尾部区域气动性能的影响,在中国空气动力研究与发展中心0.5m高超声速风洞中,在来流马赫数为5.0和6.0条件下,开展了不同落压比条件下的尾喷流干扰测压实验研究,同时采用高清纹影观测了喷流干扰区域的流场结构。实验结果表明:不同喷流落压比时,飞行器尾部区域表面压力分布差别明显,高落压比时喷流干扰作用的区域更大,压强数值更高。纹影也显示高落压比时交叉干扰激波更强、剪切层扩张更明显。喷流干扰区域已影响到了飞行器水平翼区域的压力分布,将会对飞行器操纵特性产生影响。To find out the effect of nozzle plume on aerocraft aerodynamic performance,apressure measurement experimental study on different nozzle pressure ratios（NPR）of nozzle plume was carried out at a freestream Mach number of 5.0and 6.0in the 0.5mhypersonic wind tunnel of China Aerodynamics Research and Development Center（CARDC）.The structure of the nozzle plume interaction area was obtained by using schlieren photography.The result showed obvious different pressure distribution of aerocraft afterbody area at a high NPR.When the NPR was high,the nozzle plume interaction area became bigger with a higher pressure.The schlieren pictures also showed that cross interaction shocks became stronger and the extending of the shear layer tended to be more obvious.Interaction area has affected the pressure distribution of the wings,which would affect the control ability of the aerocraft.

关 键 词：吸气式飞行器 高超声速 落压比 尾喷流 测压实验 

分 类 号：V211[航空宇航科学与技术—航空宇航推进理论与工程]