摘要
常规跨超声速风洞进行马赫数4.5试验时,常常伴有空气液化现象,造成试验数据可信度低,在高超声速风洞研制马赫数4.5喷管,具有对气流加热的能力,可以提供更加准确的试验数据。目前国内0.5 m量级高超声速风洞还不具备马赫数4.5的试验能力。通过无黏流计算方法计算轴对称喷管型面,并采用Sivells-Payne方法进行附面层修正,然后进行数值验证,证明了计算出的型面满足国军标对马赫数的设计要求,可以投入加工生产。
The Mach 4.5 test in a conventional tran-supersonic wind tunnel is often accompanied by the phenomenon of air liquefaction,resulting in low reliability of the test data.The Mach 4.5 nozzle developed in a hypersonic wind tunnel has the ability to heat the airflow,which can provide more accurate test data.At present,the test capability of Mach 4.5 is not available in China for the 0.5-meter hypersonic wind tunnel.The axisymmetric nozzle profile was calculated by the inviscid flow calculation method,and the boundary layer was modified by the Sivells-Payne method.Then,numerical simulation was carried out.It was proved that the calculated profile met the GJB design requirements of Mach number and can be put into production.
作者
黄飓
杨永能
胥继斌
杨海滨
张伟
HUANG Ju;YANG Yongneng;XU Jibin;YANG Haibin;ZHANG Wei(China Aerodynamics Research and Development Center,621000 Mianyang,China)
出处
《应用力学学报》
CAS
CSCD
北大核心
2023年第2期302-309,共8页
Chinese Journal of Applied Mechanics
关键词
高超声速
速度场
马赫数
飞行器
导弹
hypersonic speed
velocity field
Mach number
aircraft
guided missile
