Low Noise Experimental Investigation On Supersonic/Hypersonic Boundary Layer Transition And Flow Over Blunt Fins

The boundary layer of the hypersonic reentry vehicle and the airbreathing vehicle may develop to turbulent state,and the laminar flow and turbulent flow are very different in terms of friction,heat flux,separation and mixing.So far,many of the problems about transition remain unclear.Boundary layer transition is an important fundamental problem that restricts hypersonic vehicles development.Since the transition process is unsteady and non-linear,it is influenced by multi-parameter such as Mach number,Reynolds number and bluntness,and becomes a typical scientific problem on fluid mechanics to be solved;on the other hand,the transition of the boundary layer is directly related to the heat flux distribution of the hypersonic vehicle.Thus,it is of great significance for the design of the aircraft and the thermal protection.It is urgent to carry out the research on the transition and prediction.Roughness elements and protuberances are simple and effective means of boundary layer controlling.The fin is a very important airdynamic component,used to control the flight attitude of the aircraft.The shock/ shock interaction and the shock wave/ boundary layer interaction caused by the fin are also stronger,and the influence on the performance of the aircraft is affected by the interference flow caused by the interaction between the fin and the local boundary layer.Since the effect of the wind tunnel noise on the transition of boundary layer is very obvious,the conventional wind tunnel is difficult to study the transition problem.The experiments in this paper are carried out in the low noise supersonic indraft wind tunnel and low noise hypersonic wind tunnel.In this paper,Nano-tracer Planar Laser Scattering(NPLS)has been used as the major test method.Hypersonic boundary layer transition on the flat plate,hypersonic boundary layer transition on the cone,supersonic flow over circular protuberances,and hypersonic flow over blunt fins are studied in the low noise supersonic indraft wind tunnel and low noise hypersonic wind tunnel.Aiming at the application of NPLS technique in hypersonic continuous wind tunnel,the design of nano-particle generator,the research of seeding method of nano-particles,the co-operation of the particle generator and wind tunnel are carried out.After long time testing,the problems of high massflow and high pressure particle generator design,particle sedding and system co-operation have been solved.Thus,the NPLS technique can be usd successfully in the following experiments.In particular,in the low noise hypersonic Mach 6.0 wind tunnel,NPLS test technique has been used to carry out the transition test of flat boundary layer,and the transition characteristics of smooth plate and flat plate with different roughness elements have been obtained and studied.The rough element plate consists of a square rough element plate,a diamond-shaped rough element plate,a ball-roughened flat plate and a sinusoidal roughness plate.However,according to the NPLS image results,at the condition of very high Reynolds number,transition haven't been detected at the end of these flat plates.There are many reasons for the absence of transition,which may include a lower level of noise in the wind tunnel,the lack of length of the plate and the impact of the bluntness of the leading edge.The effects of Reynolds number,bluntness,roughness and angle of attack on the transition of hypersonic cone boundary layer are studied,and the fine structures of the flow field in various states are obtained.With increasing the Reynolds number,the transition point moves forward.Under the same condition,the critical Reynolds number of the blunt cone is significantly higher than that of the sharp cone,indicating that the bluntness may delay boundary layer transition.When the roughness is 50?m,the critical Reynolds number of the transition is smaller than that of the smooth cone.When the cone has the angle of attack,the boundary layer of the windward surface is more difficult to transition,and the leeward side boundary layer is eaysier to transition,compared with 0 degree of AOA.There is obvious angle of attack effect in the boundary layer transition on cones.The supersonic boundary layer transition and supersonic flow over circular protuberances have been studied by NPLS technique and WCNS-E-5 high precision numerical model.The wake characteristics of the protuberances and the influence on the transition of the boundary layer have been researched.The heights of the protuberances are namely 1mm,2mm and 4mm,the state of the boundary layer is laminar with a thickness ? of 1.2mm,so that the ratio of H/? ratio can be compared and analyzed.The fine structures of supersonic flow over protuberances have been obtained.When the height of the protuberance is close to the thickness of the local boundary layer,then the disturbance is small.The intermittent comparison of the wake flow of different protuberances has been obtained by intermittent method.In the Mach 3.8 low-noise supersonic wind tunnel,the interactions of turbulent boundary layer and the "large circular protuberances” have been studied.A total of five protuberances with different heights are studied.The diameter of the round table is 20 mm and the height is 15 mm,18mm,20 mm,25 mm and 30 mm.The separation shock wave,the shock foot and the bow detached shock wave constitute the main structure of the flow field,and the three shocks intersect at the "triple wave point".In this study,the fine structural features near the front separation zone have been obtained.It has been found that the 15 mm and 18 mm height protuberances show different shock wave structure characteristics.The height distribution characteristics of the triple point of the protuberances have been analyzed by the probabilistic statistical method.In the low noise hypersonic tunnel,the NPLS technique is used to study the flow field at the leading edge of the fins,the flow field near the fin root,and the flow field near the axle.In the experiments,the boundary layer of the surface of the aircraft is simulated by the turbulent boundary layer of the floor of the wind tunnel test section.The effects of Reynolds number,sweptback angle angle and the height of fin are studied preliminarily.As the fin axle is located inside the turbulent boundary layer of the wind tunnel test section,the complex turbulent structures are in the spanwis plane.