Experimental investigation of boundary layer transition over a delta wing at Mach number 6  

作　　者：Haibo NIU Shihe YI Xiaolin LIU Xiaoge LU Dundian GANG 

机构地区：[1]College of Aerospace Science and Engineering,National University of Defense Technology,Changsha 410073,China

出　　处：《Chinese Journal of Aeronautics》2020年第7期1889-1902,共14页中国航空学报（英文版）

基　　金：supported by the National Key Technology Research and Development Program of China(No.2016YFA0401200);the National Project for Research and Development of Major Scientific Instruments of China(No.11527802);the National Natural Science Foundation of China(Nos.91752102 and 11832018)。

摘　　要：An experimental study on the boundary layer transition over a delta wing was carried out at Mach number 6 in a quiet wind tunnel.The Nano-tracer-based Planar Laser Scattering(NPLS)and Temperature-Sensitive Paints(TSP)techniques were used to measure the fine flow field structure and the wall Stanton number of the delta wing.The influence of factors such as the angle of attack and the Reynolds number was studied.The following results were obtained.The boundary layer transition between the leading edge and the centerline was dominated by the crossflow instability.At the location of the initial appearance of the traveling crossflow waves,the Stanton number began to rise.The Stanton number reached a maximum when the crossflow waves were broken up to turbulence.Increasing the angle of attack increased the spanwise pressure gradient at the windward side of the delta wing,thereby increasing the crossflow instability and advancing the boundary layer transition front.However,increasing the angle of attack caused the transition front to move backward at the leeward side.In addition,the sensitivity of the boundary layer transition to the Reynolds number varied with the angle of attack and the region.

关 键 词：Boundary layer CROSSFLOW Delta wings HYPERSONIC Stanton number Temperature-sensitive paint Transition 

分 类 号：V211[航空宇航科学与技术—航空宇航推进理论与工程]