CAJ | 학술논문

主要介绍了一种用于高超声速湍流边界层表面参数测量的试验方法,利用边界层的表面热流值,作为判断边界层转捩与否的依据,开展边界层攻角效应转捩特性试验研究.试验在我院FD-20炮风洞中进行,马赫数为6,试验模型是半锥度5°的圆锥模型.开展了对于圆锥模型,头部半径对转捩位置的影响,小攻角情况下对圆锥模型转捩位置的影响,并且获得了在2°和3°攻角情况下圆锥模型的攻角效应转捩全局曲线.试验结果证实了,在小攻角下,随着攻角的增加,边界层迎风面转捩位置向后移动,背风面转捩位置向前移动.
주요개소료일충용우고초성속단류변계층표면삼수측량적시험방법,이용변계층적표면열류치,작위판단변계층전렬여부적의거,개전변계층공각효응전렬특성시험연구.시험재아원FD-20포풍동중진행,마혁수위6,시험모형시반추도5°적원추모형.개전료대우원추모형,두부반경대전렬위치적영향,소공각정황하대원추모형전렬위치적영향,병차획득료재2°화3°공각정황하원추모형적공각효응전렬전국곡선.시험결과증실료,재소공각하,수착공각적증가,변계층영풍면전렬위치향후이동,배풍면전렬위치향전이동.
To investment boundary layer transition characteristic with attack angle effect,an Experiment method was introduced to measure the hypersonic turbulence boundary layer parameters which used surface heat flux as basis to judge the boundary transition or not.The experiment was carried out in FD-20 piston gun tunnel at Mach number 6 using 5°haft angle circular cone model.The nose tip radius and small attack angle effects on transition position on circular cone model were researched and attack angle effect transition curves were obtained at 2°and 3°.The test results showed that at small attack angle condition with increasing of angle of attack transition position move to bottom at windward and to nose tip at leeward in the boundary layer.