This paper focused on the fundamental and applied research of turbulent flows encountered in the hypersonic flight of aerospace vehicles,which take place in the boundary layer and mixing layer.As to the plate boundary...This paper focused on the fundamental and applied research of turbulent flows encountered in the hypersonic flight of aerospace vehicles,which take place in the boundary layer and mixing layer.As to the plate boundary layer,LES approach has been used to simulate the flows over compression corners and incident shock waves,revealing that turbulent flows would significantly inhibit the boundary layer separation caused by shock wave-boundary layer interaction(SWBLI).The boundary layer transition over a circular cone has been analyzed through stability analysis and wind-tunnel test,by which the angle-of-attack effect in case of small angle of attack has been studied.Non-linear evolution process and secondary instability structure in the supersonic mixing layer(Mc=0.5) were initially figured out through the study of mixing layer,and knowledge of the flow control mechanism of the boundary layer and mixing enhancement mechanism of the mixing layer has been obtained through this research.Artificial boundary-layer transition technique based on subharmonic resonance has been proposed and applied to the flow control in a scramjet inlet,inhibiting the flow separation of the boundary layer while improving the inlet performance.To guarantee the mixing of kerosene and supersonic airflow in the scramjet combustor,the mixing enhancement method based on subharmonic resonance has been adopted and a concept of combustor with smooth wall and low internal drag has been proposed for ignition and stable combustion.Finally,future turbulence research and technological development of aerospace vehicles is predicted.展开更多
The results of experimental investigation of laminar-turbulent transition in three-dimensional flow under the high continuous pressure gradient including the flow with local boundary layer separation are presented. Th...The results of experimental investigation of laminar-turbulent transition in three-dimensional flow under the high continuous pressure gradient including the flow with local boundary layer separation are presented. The experimental studies were performed within the Mach number range from 4 to 6 and Reynolds number 10-60×106 1/m, the angles of attack were 0° and 5°. The experiments were carried out on the three-dimensional convergent inlet model with and without sidewalls. The influence of artificial tubulator of boundary layer on transition and flow structure was studied. The conducted researches have shown that adverse pressure gradient increase hastens transition and leads to decrease of transition area length. If pressure gradient rises velocity profile fullness increases and profile transformation from laminar to turbulent occurs. As a result of it the decrease of separation area length occurs. The same effect was reached with Reynolds number increase. These results are compared with the data on two-dimensional model with longitudinal curvature.展开更多
The effect of magnetohydrodynamic(MHD)plasma actuators on the control of hypersonic shock wave/turbulent boundary layer interactions is investigated here using Reynolds-averaged Navier-Stokes calculations with low mag...The effect of magnetohydrodynamic(MHD)plasma actuators on the control of hypersonic shock wave/turbulent boundary layer interactions is investigated here using Reynolds-averaged Navier-Stokes calculations with low magnetic Reynolds number approximation.A Mach 5 oblique shock/turbulent boundary layer interaction was adopted as the basic configuration in this numerical study in order to assess the effects of flow control using different combinations of magnetic field and plasma.Results show that just the thermal effect of plasma under experimental actuator parameters has no significant impact on the flow field and can therefore be neglected.On the basis of the relative position of control area and separation point,MHD control can be divided into four types and so effects and mechanisms might be different.Amongst these,D-type control leads to the largest reduction in separation length using magnetically-accelerated plasma inside an isobaric dead-air region.A novel parameter for predicting the shock wave/turbulent boundary layer interaction control based on Lorentz force acceleration is then proposed and the controllability of MHD plasma actuators under different MHD interaction parameters is studied.The results of this study will be insightful for the further design of MHD control in hypersonic vehicle inlets.展开更多
NASA rotor 37 was used as a ’blind’ test case for turbomachinery CFD by the Turbimachinery Com-mittee of the IGTI. The rotor is a transonic compressor with a tip speed of 454 m/s (15OO ft/s) anda relatively high pre...NASA rotor 37 was used as a ’blind’ test case for turbomachinery CFD by the Turbimachinery Com-mittee of the IGTI. The rotor is a transonic compressor with a tip speed of 454 m/s (15OO ft/s) anda relatively high pressure ratio of 2.1. It was tested in isolation with a circumferentially uniform inletflow so that the flow through it should be steady apart from any effects of passage to passage geometry variation and mechanical vibration. As such it represents the simplest possible type of test forthree-dimensional turbomachinery flow solvers. However, the rotor still presents a real challenge to3D viscous flow solvers because the shock wave-boundary layer interaction is strong and the effects ofviscosity are dominant in determining the flow deviation and hence the pressure ratio. Eleven ’blind’solutions were submitted and in addition a ’non-blind’ solution was used to prepare for the exercise-This paper reviews the fiow in the test case and the comparisons of the CFD solutions with the testdata. Lessons for both the Flow Physics in transonic has and for the application of CFD to suchmachines are pointed out.展开更多
Strong, normal shock wave, terminating a local supersonic area on an airfoil, not only limits aerodynamic performance but also becomes a source of a high-speed impulsive helicopter noise. The application of a passive ...Strong, normal shock wave, terminating a local supersonic area on an airfoil, not only limits aerodynamic performance but also becomes a source of a high-speed impulsive helicopter noise. The application of a passive control system （a cavity covered by a perforated plate） on a rotor blade should reduce the noise created by a moving shock. This article covers the numerical implementation of the Bohning/Doerffer transpiration law into the SPARC code and includes an extended validation against the experimental data for relatively simple geometries of transonic nozzles. It is a first step towards a full simulation of a helicopter rotor equipped with a noise reducing passive control device in hover and in forward flight conditions.展开更多
Transition prediction of the supersonic boundary layer on a cone with small angle of attack and Mach number 3.5 is investi-gated under the consideration of receptivity to slow acoustic waves, as the acoustic waves are...Transition prediction of the supersonic boundary layer on a cone with small angle of attack and Mach number 3.5 is investi-gated under the consideration of receptivity to slow acoustic waves, as the acoustic waves are the main environmental distur-bances in a conventional, i.e. non-quiet, wind tunnel. It is shown that the e-N method can still yield fairly satisfactory results incomparison with those obtained in wind tunnel experiments, provided that the boundary layer receptivity to slow acousticwaves is properly taken into account, including the dependence of the amplitude of disturbances on the frequency andstream-wise location. Neither the conventional e-N method nor the improved e-N method can yield correct result of transitionprediction, because the receptivity mechanisms considered there are not in accord with the real situation in the wind tunnel.展开更多
The present contribution describes two prediction methods for flows around transonic airfoils, including shock control devices. The wliole work was done in the frame of the European Shock Control Inves tigation Projec...The present contribution describes two prediction methods for flows around transonic airfoils, including shock control devices. The wliole work was done in the frame of the European Shock Control Inves tigation Project EUROSHOCK-AER-2, and the global objective was the improvement of the flight performance, in transonic speed, in terms of cruise speed, fuel consumption and exhaust emissions for both laminar and turbulent wings. More specilically the "passive" control of shock/boundary layer interaction, whereby part of the solid suLrfaCe of the airfoil is replaced by a porous surface over a shallow cavity, has been shown to be a means of improving the aerodynamic characteristics of supercritical airfoils.展开更多
A description and results of tests of a new small-scale gage for direct measurement of skin friction force are presented in the paper.The gage design provides separated measurement of longitudinal and transversal comp...A description and results of tests of a new small-scale gage for direct measurement of skin friction force are presented in the paper.The gage design provides separated measurement of longitudinal and transversal component of friction force.Application of this scheme provides high sensitivity and necessary high-frequency response of the gage.The tests of the gage were carried out in a blow down wind tunnel at Mach numbers of 2 and 4 within the range of Reynolds numbers Rex from 0.8 to 5 million and in the hot-shot wind tunnel at Mach number 6 and Reynolds numbers Rex=2.5-10 million.The measurements of skin friction were carried out on a flat plate and on a ramp beyond the shock wave.Simultaneously with the direct measurement of friction in the blow down wind tunnel,the measurements of profiles of average velocities and mass flow rate pulsations were realised.Analysis of measurement errors has shown that the friction gage permits to measure skin friction coefficient on a flat plate with mistake not more than 10%.展开更多
The flow in transonic diffusers as well as in supersonic air intakes becomes often unsteady due to shock wave boundary layer interaction. The oscillations may be induced by natural separation unsteadiness or may be fo...The flow in transonic diffusers as well as in supersonic air intakes becomes often unsteady due to shock wave boundary layer interaction. The oscillations may be induced by natural separation unsteadiness or may be forced by boundary conditions. Significant improvement of CFD tools, increase of computer resources as well as development of experimental methods have again.drawn the attention of researchers to this topic. To investigate the problem forced oscillations of transonic turbulent flow in asymmetric two-dimensional Laval nozzle were considered. A viscous, perfect gas flow, was numerically simulated using the Reynolds-averaged compressible Navier-Stokes solver SPARC, employing a two-equation, eddy viscosity, turbulence closure in the URANS approach.For time-dependent and stationary flow simulations, Mach numbers upstream of the shock between 1.2 and 1.4 were considered. Comparison of computed and experimental data for steady states generally gave acceptable agreement. In the case of forced oscillations, a harmonic pressure variation was prescribed at the exit plane resulting in shock wave motion. Excitation frequencies between 0 Hz and 1024 Hz were investigated at the same pressure amplitude.The main result of the work carried out is the relation between the amplitude of the shock wave motion and the excitation frequency in the investigated range. Increasing excitation frequency resulted in decreasing amplitude of the shock movement. At high frequencies a natural mode of shock oscillation (of small amplitude) was observed which is not sensitive to forced excitement.展开更多
Two types of flow configurations with bleed their aerodynamic thermal loads and related in two-dimensional hypersonic flows flow structures at choked conditions. are numerically examined to investigate One is a turbul...Two types of flow configurations with bleed their aerodynamic thermal loads and related in two-dimensional hypersonic flows flow structures at choked conditions. are numerically examined to investigate One is a turbulent boundary layer flow without shock impingement where the effects of the slot angle are discussed, and the other is shock wave boundary layer in- teractions where the effects of slot angle and slot location relative to shock impingement point are surveyed. A key separation is induced by bleed barrier shock on the upstream slot wall, resulting in a localized maximum heat flux at the reattachment point. For slanted slots, the dominating flow patterns are not much affected by the change in slot angle, but vary dramatically with slot location relative to the shock impingement point. Different flow structures are found in the case of normal slot, such as a flow pattern similar to typical Laval nozzle flow, the largest separation bubble which is almost independent of the shock position. Its larger detached distance results in 20% lower stagnation heat flux on the downstream slot corner, but with much wider area suffering from severe thermal loads. In spite of the complexity of the flow patterns, it is clearly revealed that the heat flux generally rises with the slot location moving downstream, and an increase in slot angle from 20° to 40° reduces 50% the heat flux peak at the reattachment point in the slot passage. The results further indicate that the bleed does not raise the heat flux around the slot for all cases except for the area around the downstream slot corner. Among all bleed configurations, the slot angle of 40° located slightly upstream of the incident shock is regarded as the best.展开更多
文摘This paper focused on the fundamental and applied research of turbulent flows encountered in the hypersonic flight of aerospace vehicles,which take place in the boundary layer and mixing layer.As to the plate boundary layer,LES approach has been used to simulate the flows over compression corners and incident shock waves,revealing that turbulent flows would significantly inhibit the boundary layer separation caused by shock wave-boundary layer interaction(SWBLI).The boundary layer transition over a circular cone has been analyzed through stability analysis and wind-tunnel test,by which the angle-of-attack effect in case of small angle of attack has been studied.Non-linear evolution process and secondary instability structure in the supersonic mixing layer(Mc=0.5) were initially figured out through the study of mixing layer,and knowledge of the flow control mechanism of the boundary layer and mixing enhancement mechanism of the mixing layer has been obtained through this research.Artificial boundary-layer transition technique based on subharmonic resonance has been proposed and applied to the flow control in a scramjet inlet,inhibiting the flow separation of the boundary layer while improving the inlet performance.To guarantee the mixing of kerosene and supersonic airflow in the scramjet combustor,the mixing enhancement method based on subharmonic resonance has been adopted and a concept of combustor with smooth wall and low internal drag has been proposed for ignition and stable combustion.Finally,future turbulence research and technological development of aerospace vehicles is predicted.
文摘The results of experimental investigation of laminar-turbulent transition in three-dimensional flow under the high continuous pressure gradient including the flow with local boundary layer separation are presented. The experimental studies were performed within the Mach number range from 4 to 6 and Reynolds number 10-60×106 1/m, the angles of attack were 0° and 5°. The experiments were carried out on the three-dimensional convergent inlet model with and without sidewalls. The influence of artificial tubulator of boundary layer on transition and flow structure was studied. The conducted researches have shown that adverse pressure gradient increase hastens transition and leads to decrease of transition area length. If pressure gradient rises velocity profile fullness increases and profile transformation from laminar to turbulent occurs. As a result of it the decrease of separation area length occurs. The same effect was reached with Reynolds number increase. These results are compared with the data on two-dimensional model with longitudinal curvature.
基金Project supported by the National Key R&D Program of China(Nos.2019YFA0405300 and 2019YFA0405203)the Chinese Scholarship Council(CSC)(No.201903170195)。
文摘The effect of magnetohydrodynamic(MHD)plasma actuators on the control of hypersonic shock wave/turbulent boundary layer interactions is investigated here using Reynolds-averaged Navier-Stokes calculations with low magnetic Reynolds number approximation.A Mach 5 oblique shock/turbulent boundary layer interaction was adopted as the basic configuration in this numerical study in order to assess the effects of flow control using different combinations of magnetic field and plasma.Results show that just the thermal effect of plasma under experimental actuator parameters has no significant impact on the flow field and can therefore be neglected.On the basis of the relative position of control area and separation point,MHD control can be divided into four types and so effects and mechanisms might be different.Amongst these,D-type control leads to the largest reduction in separation length using magnetically-accelerated plasma inside an isobaric dead-air region.A novel parameter for predicting the shock wave/turbulent boundary layer interaction control based on Lorentz force acceleration is then proposed and the controllability of MHD plasma actuators under different MHD interaction parameters is studied.The results of this study will be insightful for the further design of MHD control in hypersonic vehicle inlets.
文摘NASA rotor 37 was used as a ’blind’ test case for turbomachinery CFD by the Turbimachinery Com-mittee of the IGTI. The rotor is a transonic compressor with a tip speed of 454 m/s (15OO ft/s) anda relatively high pressure ratio of 2.1. It was tested in isolation with a circumferentially uniform inletflow so that the flow through it should be steady apart from any effects of passage to passage geometry variation and mechanical vibration. As such it represents the simplest possible type of test forthree-dimensional turbomachinery flow solvers. However, the rotor still presents a real challenge to3D viscous flow solvers because the shock wave-boundary layer interaction is strong and the effects ofviscosity are dominant in determining the flow deviation and hence the pressure ratio. Eleven ’blind’solutions were submitted and in addition a ’non-blind’ solution was used to prepare for the exercise-This paper reviews the fiow in the test case and the comparisons of the CFD solutions with the testdata. Lessons for both the Flow Physics in transonic has and for the application of CFD to suchmachines are pointed out.
文摘Strong, normal shock wave, terminating a local supersonic area on an airfoil, not only limits aerodynamic performance but also becomes a source of a high-speed impulsive helicopter noise. The application of a passive control system （a cavity covered by a perforated plate） on a rotor blade should reduce the noise created by a moving shock. This article covers the numerical implementation of the Bohning/Doerffer transpiration law into the SPARC code and includes an extended validation against the experimental data for relatively simple geometries of transonic nozzles. It is a first step towards a full simulation of a helicopter rotor equipped with a noise reducing passive control device in hover and in forward flight conditions.
基金supported by the National Natural Science Foundation of China (Grant Nos. 10632050 and 11002098)the National Basic Research Program of China (Grant No. 2009CB724103)the Specialized Research Fund for the Doctoral Program of Higher Education
文摘Transition prediction of the supersonic boundary layer on a cone with small angle of attack and Mach number 3.5 is investi-gated under the consideration of receptivity to slow acoustic waves, as the acoustic waves are the main environmental distur-bances in a conventional, i.e. non-quiet, wind tunnel. It is shown that the e-N method can still yield fairly satisfactory results incomparison with those obtained in wind tunnel experiments, provided that the boundary layer receptivity to slow acousticwaves is properly taken into account, including the dependence of the amplitude of disturbances on the frequency andstream-wise location. Neither the conventional e-N method nor the improved e-N method can yield correct result of transitionprediction, because the receptivity mechanisms considered there are not in accord with the real situation in the wind tunnel.
文摘The present contribution describes two prediction methods for flows around transonic airfoils, including shock control devices. The wliole work was done in the frame of the European Shock Control Inves tigation Project EUROSHOCK-AER-2, and the global objective was the improvement of the flight performance, in transonic speed, in terms of cruise speed, fuel consumption and exhaust emissions for both laminar and turbulent wings. More specilically the "passive" control of shock/boundary layer interaction, whereby part of the solid suLrfaCe of the airfoil is replaced by a porous surface over a shallow cavity, has been shown to be a means of improving the aerodynamic characteristics of supercritical airfoils.
文摘A description and results of tests of a new small-scale gage for direct measurement of skin friction force are presented in the paper.The gage design provides separated measurement of longitudinal and transversal component of friction force.Application of this scheme provides high sensitivity and necessary high-frequency response of the gage.The tests of the gage were carried out in a blow down wind tunnel at Mach numbers of 2 and 4 within the range of Reynolds numbers Rex from 0.8 to 5 million and in the hot-shot wind tunnel at Mach number 6 and Reynolds numbers Rex=2.5-10 million.The measurements of skin friction were carried out on a flat plate and on a ramp beyond the shock wave.Simultaneously with the direct measurement of friction in the blow down wind tunnel,the measurements of profiles of average velocities and mass flow rate pulsations were realised.Analysis of measurement errors has shown that the friction gage permits to measure skin friction coefficient on a flat plate with mistake not more than 10%.
文摘The flow in transonic diffusers as well as in supersonic air intakes becomes often unsteady due to shock wave boundary layer interaction. The oscillations may be induced by natural separation unsteadiness or may be forced by boundary conditions. Significant improvement of CFD tools, increase of computer resources as well as development of experimental methods have again.drawn the attention of researchers to this topic. To investigate the problem forced oscillations of transonic turbulent flow in asymmetric two-dimensional Laval nozzle were considered. A viscous, perfect gas flow, was numerically simulated using the Reynolds-averaged compressible Navier-Stokes solver SPARC, employing a two-equation, eddy viscosity, turbulence closure in the URANS approach.For time-dependent and stationary flow simulations, Mach numbers upstream of the shock between 1.2 and 1.4 were considered. Comparison of computed and experimental data for steady states generally gave acceptable agreement. In the case of forced oscillations, a harmonic pressure variation was prescribed at the exit plane resulting in shock wave motion. Excitation frequencies between 0 Hz and 1024 Hz were investigated at the same pressure amplitude.The main result of the work carried out is the relation between the amplitude of the shock wave motion and the excitation frequency in the investigated range. Increasing excitation frequency resulted in decreasing amplitude of the shock movement. At high frequencies a natural mode of shock oscillation (of small amplitude) was observed which is not sensitive to forced excitement.
基金supported by the National Natural Science Foundation of China(Grant Nos.91216115 and 11472279)
文摘Two types of flow configurations with bleed their aerodynamic thermal loads and related in two-dimensional hypersonic flows flow structures at choked conditions. are numerically examined to investigate One is a turbulent boundary layer flow without shock impingement where the effects of the slot angle are discussed, and the other is shock wave boundary layer in- teractions where the effects of slot angle and slot location relative to shock impingement point are surveyed. A key separation is induced by bleed barrier shock on the upstream slot wall, resulting in a localized maximum heat flux at the reattachment point. For slanted slots, the dominating flow patterns are not much affected by the change in slot angle, but vary dramatically with slot location relative to the shock impingement point. Different flow structures are found in the case of normal slot, such as a flow pattern similar to typical Laval nozzle flow, the largest separation bubble which is almost independent of the shock position. Its larger detached distance results in 20% lower stagnation heat flux on the downstream slot corner, but with much wider area suffering from severe thermal loads. In spite of the complexity of the flow patterns, it is clearly revealed that the heat flux generally rises with the slot location moving downstream, and an increase in slot angle from 20° to 40° reduces 50% the heat flux peak at the reattachment point in the slot passage. The results further indicate that the bleed does not raise the heat flux around the slot for all cases except for the area around the downstream slot corner. Among all bleed configurations, the slot angle of 40° located slightly upstream of the incident shock is regarded as the best.
