Experimental Investigation Of Hypersonic Boundary Layer Transition And Shock Wave/Turbulent Boundary Layer Interactions

Complex flow phenomena such as hypersonic boundary layer transition and shock wave and turbulent boundary layer interactions are commonly found in the high-speed flight of hypersonic vehicles,which are hot and difficult problems in the field of hypersonic aerodynamics.It is of great theoretical significance and engineering application value to accurately predict the boundary layer transition position and control the transition,and to understand the unsteady characteristics of the shock wave and turbulent boundary layer interactions,so as to effectively reduce the drag of hypersonic vehicles,optimize the design of thermal protection and improve the performance of vehicles.In this paper,with the Nano-tracer-based Planar Laser Scattering（NPLS）technique,Temperature Sensitive Paint（TSP）technique and the high-frequency fluctuation pressure measurement technique as the main technical means,the typical flow fields of hypersonic vehicles such as the hypersonic boundary layer transition of a cone with a half cone angle of 5°,the hypersonic boundary layer transition on swept flat plate and shock wave and turbulent boundary layer interactions,in-depth experimental studies were carried out in hypersonic quiet wind tunnel,hypersonic low-noise wind tunnel and supersonic low-noise wind tunnel,respectively.There is quite different between the data of the experimental environment of wind tunnel and the real flight environment of hypersonic vehicles,the noise reduction technology in the settling chamber of hypersonic quiet wind tunnel and the optimization design method of new quiet nozzle were put forward in this paper.So as to realize the quiet wind tunnel with the quiet mode（the incoming turbulence fluctuation levels is less than 0.1%）,low noise mode（the incoming turbulence fluctuation levels is 0.1～0.5%）and high noise mode（the incoming turbulence fluctuation levels is greater than 2%）.And the effective running time of the hypersonic quiet tunnel is more than 30s.The application of relevant technologies in hypersonic and supersonic direct-connected wind tunnels was carried out,and the turbulence degree of free incoming flow in the two wind tunnels is less than 0.5%,thus achieving the low noise experimental flow field condition.At the same time,based on the NPLS technique,the NPLS system for studying the time evolution process of unsteady flow structures was developed.And for the first time,eight frames of temporal-correlated fine structure images of transient flow field with the shortest time interval of 1μs was realized under experimental conditions.The highest sampling frame rate of the system reached 1MHz,and the system was successfully applied to the experiment of shock wave and turbulent boundary layer interactions.In this paper,the experimental study on the transition of hypersonic conical boundary layer with half cone angle of 5°was performed in a hypersonic quiet wind tunnel,and the effects of leading-edge bluntness,model angle of attack and the unit Reynolds number on the transition of conical boundary layer were studied,respectively.Fine flow structure images of boundary layer transition on windward and leeward sides of smooth cone were obtained by NPLS technique,and the clear"rope-like"second modal wave structures were observed under the conditions of partial unit Reynolds number(0.27×10⁷m^-1～1.30×10⁷m^-1)and small angle of attack（angles of attack of0°～5°）.Under the same angle of attack,with the increase of unit Reynolds number,the boundary layer transition position on the windward and leeward sides of the cone is obviously advanced.When the angle of attack increases to 5°,the boundary layer transition position on the leeward side is very close to the leading edge due to the influence of cross flow.Under the same or similar Reynolds number,with the increase of angle of attack,the boundary layer transition position on the windward side of the conical model is obviously delayed,while the boundary layer transition position on the leeward side is advanced.In a certain range of leading-edge bluntness（R=0.1～7.0mm）,with the increase of leading-edge bluntness,the boundary layer transition positions on windward and leeward sides of conical model are obviously delayed.Under the condition of the angle of attack was 0°and the unit Reynolds number of 1.0×10⁷m^-1,through the analysis of the time evolution process of the second modal wave structure in two images with time correlation（Δt=10μs）,it is found that the statistical wavelength is5.39mm,and the propagation velocity of the second modal wave along the flow direction is 747m/s,in which the propagation velocity of the second mode wave is about0.87 times the mainstream velocity.The characteristic frequency corresponding to the second modal wave is 138.6k Hz.The development law of the second mode wave in the conical boundary layer is analyzed by using the high-frequency pulsating pressure sensor.The characteristic frequency of the second modal wave decreases gradually during the development along the flow direction,and the amplitude of the disturbance wave increases rapidly at first and then gradually decays to disappear.Under the condition of angle of attack is 0°,with the increase of Reynolds number,the characteristic frequency of the second mode wave gradually increases,and the starting position of the disturbance wave gradually moves towards the leading edge of the cone.The scatter distribution of the disturbance wave before and after the experimental unit Reynolds number of 1.39×10⁷m^-1 is antisymmetric relative to the linear fitting trend line.With the increase of the angle of attack,the position of the second mode wave on the windward and leeward sides start to increase obviously.Due to the influence of cross flow,the second mode wave signal could not be detected at most measuring points at angle of attack of 5°.Compared with the measurement results obtained by NPLS technique and high-frequency pulsating pressure measurement technique,the position of the second mode wave signal measured by the sensor is slightly earlier than the spatial position of the flow structure distribution.But the NPLS technique can more intuitively reflect the changes of the flow structure in the process of boundary layer development,and can directly judge the interval range of boundary layer transition through image statistical analysis.The effects of Reynolds number and model angle of attack on boundary layer transition of a smooth swept plate with blunt leading-edge were studied in hypersonic low-noise wind tunnel.The instantaneous fine structures of the boundary layer on the swept flat plate in streamwise and spanwise planes have been investigated based on the NPLS technique,respectively.The spatiotemporal evolution characteristics of the boundary layer translating from laminar to turbulence were analyzed.With the increase of Reynolds number,the position boundary layer transition on windward side was advanced.In the experimental Reynolds number range(1.0×10⁷m^-1～2.6×10⁷m^-1),the parallel distance between boundary layer transition position and swept leading edge is less than 120mm.Cross-flow waves in swept plate flow structures were studied.And cross-flow waves in spanwise images with Reynolds numbers in the range of1.0×10⁷m^-1～2.1×10⁷m^-1 at angle of attack of 0°were analyzed.The results show that the wavelength range of cross-flow waves is 4.05mm～4.75mm,and the characteristic frequency range is 19.7k Hz～111.2k Hz.With the increase of Reynolds number,the characteristic wavelength of cross-flow waves become longer,and the characteristic frequency of cross-flow waves become higher.Using TSP technique,the temperature field distribution during the development of the wall boundary layer was obtained,and the basic transition front of the flat boundary layer with smooth swept blunt leading edge was analyzed,and its change rule is consistent with the NPLS results.However,the use of high-frequency pulsating pressure sensors did not measure the useful characteristic signals about the cross-flow waves.In this paper,the experimental study on the unsteady characteristics of interactions between incident shock waves generated by shock generators with different deflection angles and turbulent boundary layer was performed in a Mach 3.4 supersonic low-noise wind tunnel at the unit Reynolds number of 6.30×10⁶/m^-1.For the first time,the high-definition fine flow structure images of transient flow field with SWTBLI process with eight frames of temporal-correlated were obtained under experimental conditions.It is preliminarily obtained that the time scale of motion characteristics of large-scale structures,such as separation bubbles,is greater than 150μs,which is about 8.3δ/U_∞of mainstream.And the characteristic frequency of low-frequency motion of SWTBLI is less than 6.8k Hz.At the same time,the influence of boundary layer thickness on SWTBLI was analyzed.The results show that with the increase of the incoming boundary layer thickness,the group velocity in the development process of vortex structure in the turbulent boundary layer does not change significantly.As the thickness of the boundary layer entering the separation bubble increases,the overall growth height of the separation bubble also increases.The statistics based on the images of NPLS flow visualization,it is found that the oscillation position of induced shock wave satisfies normal distribution under different incident shock intensity.Using TSP technique,the wall temperature distribution of model of interactions region under different incident shock wave conditions was obtained.Based on the unsteady characteristics of SWTBLI measured by pulsating pressure under different deflection angles,it is found that the frequency distribution in the interactions area is dominated by low-frequency signals below 4k Hz.After the intensity of incident shock wave increases（θ=10°～20°）,the low-frequency oscillation of separation bubble has a peak frequency signal at500Hz～1000Hz.