Research On Flow Mechanism Of Hypersonic Plate/rudder Interference And Three Dimensional Boundary Layer Transition

As a typical component installed on the hypersonic maneuvering aircraft,the rudder plays an important role in controlling the attitude and improving the maneuverability of the aircraft.However,in the hypersonic flow,the complicated flow phenomena induced by the rudder such as shock wave-shock wave interference,shock wave-boundary layer interaction,boundary layer separation and reattachment,etc.,cause severe local aerodynamic heating.Especially for the rudder with the installation gap,the existence of the gap and rudder shaft further increases the complexity of the flow structures and the aerodynamic heating,which requires special attention.On the other hand,to improve the range and maneuverability,swept wings and scramjet engines have become important components used in hypersonic aircraft.Both the hypersonic swept wing and the forebody boundary layer are complex three-dimensional boundary layers,and their laminar-turbulent transition has significant effects on the aerodynamic heating,surface friction and separation,the engine inlet’s starting characteristics,fuel mixing,etc.The research on three dimensional boundary layer transition of hypersonic swept wing and forebody has important guiding significance for the design of hypersonic aircraft.Aiming at the swept rudder-induced flow field and the three-dimensional boundary layer transition under hypersonic conditions,a series of experiments were conducted in the hypersonic low-noise wind tunnel using the Nano-tracer-based Planar Laser Scattering(NPLS)and Temperature Sensitive Paint(TSP).To assist the analysis of experimental results,the corresponding numerical simulation analysis were carried out.The combination of the experimental results and the numerical simulation has played a good role in the analysis of the flow mechanism of the hypersonic flow.Due to the complexity of the flow structures of the hypersonic plate/rudder interference and the three-dimensional boundary layer transition,especially for the region near the gap,rudder shaft,leading edge and other small details affecting the flow structures,the visualization of the flow in the plane perpendicular to the flow direction is very necessary.To overcome the difficulty of visualizing the flow structures in the plane perpendicular to the flow direction,research on the methods of shooting and image correction was carried out.Referring to the 3D-PIV(stereoscopic Particle Image Velocimetry),the off-axis imaging layout of Scheimpflüg was applied to take off-axis shooting of the flow field with a single CCD.The image geometric correction is carried out with checkerboard calibrations and the second-order polynomial distortion fitting function.Cubic interpolation is used for Gray-scale interpolation,which better maintains the flow details while ensuring the accuracy of correction.The rudder gap has a very important influence on the structures of the hypersonic plate/rudder interference flow field.Thus,the research on the hypersonic plate/rudder interference flow field is divided into two chapters: the rudder without gap and the rudder with gap.The interference flow field without gap is mainly characterized by the thin boundary layer on both sides of the rudder and the horseshoe separation vortex induced by the bow-shaped detached shock wave.The heat flux distribution is closely related to the flow structures.The vortex,reattachment lines,and thin boundary layer regions usually correspond to higher levels of heat flux.Unlike the state without rudder gap,the hypersonic plate/rudder interference flow field with rudder gap is more complicated.The ratio of the gap height to the thickness of the incoming boundary layer(h/ δ)becomes an important factor affecting the characteristics of the flow field.When h/ δ is less than a certain critical value,the rudder root downwash flow can still maintain a high reattachment strength on the plate near the rudder leading edge.Thus,the flow characteristics are similar to the case without rudder gap.When h/ δ is larger,the reattachment intensity of the rudder root downwash flow is obviously weakened,and the corresponding reattachment line diverges earlier,forming a high heat flow zone with relatively uniform magnitude.For the separation zone near the shaft,the increase of the gap height increases the separation strength and the size of the separation zone.The high heat flux area along the reattachment line of the rudder shaft separation zone becomes the region with the peak heat flux in the flow field.The NPLS flow visualization results show that the boundary layer transition of the hypersonic swept wing surface can be divided into two regions,namely,the transition dominated by the wing root separation vortex at the wing root and the crossflow dominated transition on the wing surface at the far wing root.Furthermore,based on the heat flux distribution obtained by the TSP experiment,the front edge of the boundary layer transition array is zigzag,along with many streamwise heat flux strips.According to these phenomena,it can be inferred that the crossflow vortices causing the transition are stationary vortices.The crossflow vortex is generated near the leading edge,and its distribution is basically along the direction of the wall streamline.The instability position of the crossflow vortex is consistent with the position of the zigzag transition front of the high heat flux zone.The increase in Reynolds number and the angle of attack have similar effects.That is,the boundary layer becomes thinner and the transition position moves forward.Meanwhile,the heat flux caused by transition can be greatly increased.The study of the transition of the hypersonic forebody boundary layer found that the boundary layer of the forebody centerline thickened significantly due to the lateral pressure gradient from both sides to the centerline,and its flow evolution was more stable.The transition of the boundary layer on both sides of the centerline is dominated by the crossflow strips,which are generated near the front edge,and are basically distributed along the direction of the surface streamline near the leading edge.With the increase of the angle of attack,the thickness of the boundary layer on the forebody becomes thinner,the inclination angle of the crossflow strips decreases,and the instability position moves upstream,so that the transition heat flux region with a“pulmonary lobe shape” extends upstream with the heat flux amplitude increasing significantly.