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)...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.展开更多
In order to apply the air fin successfully and ensure the maneuverability of hypersonic vehicle, a key problem to be studied urgently is the heat flux brought by the fin mounting gap.The appearance of mounting gap and...In order to apply the air fin successfully and ensure the maneuverability of hypersonic vehicle, a key problem to be studied urgently is the heat flux brought by the fin mounting gap.The appearance of mounting gap and fin shaft can induce many complex flow structures which need more attentions to be investigated. Under Ma 6, Nano-tracer-based Planar Laser Scattering(NPLS)and Temperature Sensitive Paints(TSP) were applied to visualize and measure transient flow structures and heat flux distribution of a swept fin-induced flow field with different height mounting gaps. Complementarily, Reynolds-averaged N-S equations were solved with k-x SST turbulent model. The heat flux distribution results of numerical simulation and TSP observed the change of high heat flux region with different mounting gap, both in position and magnitude. The streamlines based on Computational Fluid Dynamics(CFD) and flow visualization results obtained by NPLS revealed the cause of high heat flux region. The high heat flux region in this flow field is mainly related to the reattachment of vortex and flow stagnation. The increase of gap height can lead to stronger gap overflow and shaft-induced horseshoe vortex, which are source of the high heat flux around the fin. The case with the highest mounting gap(4 mm) en-counters the most severe aerodynamic heating, both on the surface of fin and plate. Thus, under the premise of ensuring the flexibility of the fin, the gap should be set as small as possible.展开更多
The experiment is conducted to investigate the effect of expansion on the shock wave boundary layer interaction near a compression ramp. The small-angle expansion with an angle degree of 5° occurs at different po...The experiment is conducted to investigate the effect of expansion on the shock wave boundary layer interaction near a compression ramp. The small-angle expansion with an angle degree of 5° occurs at different positions in front of the compression ramp. The particle image velocimetry and flow visualization technology show the flow structures, velocity field, and velocity fluctuation near the compression ramp. The mean pressure distribution, pressure fluctuation, and power spectral density are measured by high-frequency response pressure transducers. The experimental results indicate that the expansion before the compression ramp position affects the shock wave boundary layer interaction to induce a large-scale separation. But the velocity fluctuation and pressure fluctuation are attenuated near the large-scale flow separation region. When the expansion occurs closer to the compression ramp, the expansion has a more significant impact on the flow. The fluctuation of velocity and pressure is significantly attenuated, and the wall pressure rise of the separation point is reduced obviously. And the characteristic low-frequency spectrum signal related to the unsteadiness of the shock wave boundary layer interaction is significantly suppressed. In addition, variation of the separation region scale at different compression angle degrees is distinctive with the effect of expansion.展开更多
An experimental study was conducted on turbulent separation behaviors induced by blunt fins with different sweep angles at Mach number 6.0.The Nano-particle based Planar Laser Scattering technique(NPLS)was applied to ...An experimental study was conducted on turbulent separation behaviors induced by blunt fins with different sweep angles at Mach number 6.0.The Nano-particle based Planar Laser Scattering technique(NPLS)was applied to visualize the flowfield,complemented by pressure tests.Sweep angles of the fins were 10°,20°,...,60°,with the same leading edge diameter of 10 mm.Fine structures of the interference flowfield induced by blunt fins have been obtained,including the shock systems and vortexes.It was found that the features and shapes of the detached shock depended on sweep angle.When sweep angle<50°,the detached shock appeared as the form of trailing shock,and the supersonic jet with its reflection could be observed.The detached shock would be curved for the 50°and 60°fins and became a transmitted shock.The Scale-Invariant Feature Transform(SIFT)was successfully applied to obtain the velocity field from NPLS images,and the extent of the separated region was found to decrease with increasing sweep angle.No separation appeared as sweep angle>30°.Two peak values were detected on the centerline pressure distribution.The first peak did not rely on sweep angle,while the second peak value decreased with increasing sweep angle.展开更多
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)。
文摘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.
基金supported by the National Key Technology Research and Development Program of China (No. 2019YFA0405300)the National Project for Research and Development of Major Scientific Instruments of China (No. 11527802)the National Natural Science Foundation of China (No. 11832018)。
文摘In order to apply the air fin successfully and ensure the maneuverability of hypersonic vehicle, a key problem to be studied urgently is the heat flux brought by the fin mounting gap.The appearance of mounting gap and fin shaft can induce many complex flow structures which need more attentions to be investigated. Under Ma 6, Nano-tracer-based Planar Laser Scattering(NPLS)and Temperature Sensitive Paints(TSP) were applied to visualize and measure transient flow structures and heat flux distribution of a swept fin-induced flow field with different height mounting gaps. Complementarily, Reynolds-averaged N-S equations were solved with k-x SST turbulent model. The heat flux distribution results of numerical simulation and TSP observed the change of high heat flux region with different mounting gap, both in position and magnitude. The streamlines based on Computational Fluid Dynamics(CFD) and flow visualization results obtained by NPLS revealed the cause of high heat flux region. The high heat flux region in this flow field is mainly related to the reattachment of vortex and flow stagnation. The increase of gap height can lead to stronger gap overflow and shaft-induced horseshoe vortex, which are source of the high heat flux around the fin. The case with the highest mounting gap(4 mm) en-counters the most severe aerodynamic heating, both on the surface of fin and plate. Thus, under the premise of ensuring the flexibility of the fin, the gap should be set as small as possible.
基金supported by the National Key Technology Research and Development Program of China(No.2019YFA0405300)the National Natural Science Foundation of China(No.11902354).
文摘The experiment is conducted to investigate the effect of expansion on the shock wave boundary layer interaction near a compression ramp. The small-angle expansion with an angle degree of 5° occurs at different positions in front of the compression ramp. The particle image velocimetry and flow visualization technology show the flow structures, velocity field, and velocity fluctuation near the compression ramp. The mean pressure distribution, pressure fluctuation, and power spectral density are measured by high-frequency response pressure transducers. The experimental results indicate that the expansion before the compression ramp position affects the shock wave boundary layer interaction to induce a large-scale separation. But the velocity fluctuation and pressure fluctuation are attenuated near the large-scale flow separation region. When the expansion occurs closer to the compression ramp, the expansion has a more significant impact on the flow. The fluctuation of velocity and pressure is significantly attenuated, and the wall pressure rise of the separation point is reduced obviously. And the characteristic low-frequency spectrum signal related to the unsteadiness of the shock wave boundary layer interaction is significantly suppressed. In addition, variation of the separation region scale at different compression angle degrees is distinctive with the effect of expansion.
基金supported by the National Natural Science Foundation of China(No.11902354)。
文摘An experimental study was conducted on turbulent separation behaviors induced by blunt fins with different sweep angles at Mach number 6.0.The Nano-particle based Planar Laser Scattering technique(NPLS)was applied to visualize the flowfield,complemented by pressure tests.Sweep angles of the fins were 10°,20°,...,60°,with the same leading edge diameter of 10 mm.Fine structures of the interference flowfield induced by blunt fins have been obtained,including the shock systems and vortexes.It was found that the features and shapes of the detached shock depended on sweep angle.When sweep angle<50°,the detached shock appeared as the form of trailing shock,and the supersonic jet with its reflection could be observed.The detached shock would be curved for the 50°and 60°fins and became a transmitted shock.The Scale-Invariant Feature Transform(SIFT)was successfully applied to obtain the velocity field from NPLS images,and the extent of the separated region was found to decrease with increasing sweep angle.No separation appeared as sweep angle>30°.Two peak values were detected on the centerline pressure distribution.The first peak did not rely on sweep angle,while the second peak value decreased with increasing sweep angle.