摘要
本文描述了二元跨音扩压器中激波附面层相互作用区的流场。这种相互作用现象常出现在超音速飞机进气道或冲压发动机中,并可能会严重影响发动机的稳定工作。实验所用模型扩压角为6°,面积比为1.6。研究了直激波波前M数为1.47时的相互作用区流场,并与同等激波强度下的平板附面层流场作了比较,同时初步比较了等压出口和奇速出口两种边界条件下流场的异同。实验结果表明:在激波下游,分离包的尺度以及耗散层的增长都比平板情况要大许多倍。λ波后存在着一系列直激波。迅速增长的耗散层与来自λ激波分叉点的滑流层约在波后100mm处汇合。出口边界的变化对激波振荡的特征频率和分离包内的紊流度有较大影响,而对激波振荡的RMS值影响甚微。
This paper presents an experimental investigation of the shock-wave boundary-layer interaction in a twp-dimensional transonic diffuser. The diffuser has a divergence angle of 6° and an area ratio of 1.6.The Mach number in front of the shock is 1.47 and that at the foot of the λ shock is 1.42. The adverse pressure gradient downstream of the shock prevents the separated boundary layer from reattaching, and this results in a very large separation bubble. Downstream of the main shock, there is a series of normal shock waves. The development of the dissipative layer is very rapid and it soon merges with the slip stream emanating from the bifurcation point of the λ shock.The test results obtained under constant pressure exit condition are compared with those obtained under sonic throat exit condition. The differences are mainly in characteristic frequencies of shock oscillation and in turbulence levels inside the separation bubbles. RMS of the shock oscillation remains nearly the same for both exit conditions.
出处
《南京航空航天大学学报》
EI
CAS
1985年第2期56-69,共14页
Journal of Nanjing University of Aeronautics & Astronautics
