As a multidisciplinary phenomenon,panel aeroelasticity in shock-dominated flow is featured by two primary interactions:Fluid-Structure Interactions(FSIs)and Shock-Boundary Layer Interactions(SBLIs).The former raises s...As a multidisciplinary phenomenon,panel aeroelasticity in shock-dominated flow is featured by two primary interactions:Fluid-Structure Interactions(FSIs)and Shock-Boundary Layer Interactions(SBLIs).The former raises structural concerns,and the latter is of aerodynamic interest.Thus,panel aeroelasticity in shock-dominated flow represents a vital topic for the development and optimization of supersonic vehicles and propulsion systems.This review systematically summarizes recent advances in the methodologies applied to capture structural and fluid dynamics,including theoretical models,numerical simulations,and wind tunnel experiments.The application of data-driven modal decomposition,an advanced technique to extract physically crucial features,on the topic is introduced.From the perspective of FSIs,the distinctive aeroelastic behaviors in shock-dominated flow,including hysteresis phenomena and nonlinear responses,are highlighted.From the perspective of SBLIs,the modifications in their spatial and temporal characteristics imposed by the aeroelastic responses are emphasized.Motivated by the interaction between the shock waves and structural response,different strategies have been proposed to implement aeroelastic suppression and shock control,which have the potential to enhance structural safety and aerodynamic performance in the next generation of high-speed flight vehicles.展开更多
For hypersonic air-breathing vehicles,the V-shaped leading edges(VSLEs)of supersonic combustion ramjet(scramjet)inlets experience complex shock interactions and intense aerodynamic loads.This paper provides a comprehe...For hypersonic air-breathing vehicles,the V-shaped leading edges(VSLEs)of supersonic combustion ramjet(scramjet)inlets experience complex shock interactions and intense aerodynamic loads.This paper provides a comprehensive review of flow characteristics at the crotch of VSLEs,with particular focus on the transition of shock interaction types and the variation of wall heat flux under different freestream Mach numbers and geometric configurations.The mechanisms governing shock transition,unsteady oscillations,hysteresis,and three-dimensional effects in VSLE flows are first examined.Subsequently,thermal protection strategies aimed at mitigating extreme heating loads are reviewed,emphasizing their relevance to practical engineering applications.Special attention is given to recent studies addressing thermochemical nonequilibrium effects on VSLE shock interactions,and the limitations of current research are critically assessed.Finally,perspectives for future investigations into hypersonic VSLE shock interactions are outlined,highlighting opportunities for advancing design and thermal management strategies.展开更多
The stability of supersonic inlets faces challenges due to various changes in flight conditions,and flow control methods that address shock wave/boundary layer interactions under only one set of conditions cannot meet...The stability of supersonic inlets faces challenges due to various changes in flight conditions,and flow control methods that address shock wave/boundary layer interactions under only one set of conditions cannot meet developmental requirements.This paper proposes an adaptive bump control scheme and employs dynamic mesh technology for numerical simulation to investigate the unsteady control effects of adaptive bumps.The obtained results indicate that the use of moving bumps to control shock wave/boundary layer interactions is feasible.The adaptive control effects of five different bump speeds are evaluated.Within the range of bump speeds studied,the analysis of the flow field structure reveals the patterns of change in the separation zone area during the control process,as well as the relationship between the bump motion speed and the control effect on the separation zone.It is concluded that the moving bump endows the boundary layer with additional energy.展开更多
Accurate predictions of Shock Waves and Boundary Layer Interaction(SWBLI)and strong Shock Waves and Wake Vortices Interaction(SWWVI)in a highly-loaded turbine propose challenges to the currently widely used Reynolds-A...Accurate predictions of Shock Waves and Boundary Layer Interaction(SWBLI)and strong Shock Waves and Wake Vortices Interaction(SWWVI)in a highly-loaded turbine propose challenges to the currently widely used Reynolds-Averaged Navier-Stokes(RANS)model.In this work,the SWBLI and the SWWVI in a highly-loaded Nozzle Guide Vane(NGV)are studied using a hybrid RANS/LES strategy.The Turbulence Kinetic Energy(TKE)budget and the Proper Orthogonal Decomposition(POD)method are used to analyze flow mechanisms.Results show that this hybrid RANS/LES method can obtain detailed flow structures for flow mechanisms analysis.Strong shock waves induce boundary layer separation,while the presence of a separation bubble can in turn lead to a Mach reflection phenomenon.The shock waves cause trailing-edge vortices to break clearly,and the wakes,in turn,can change the shocks intensity and direction.Furthermore,the Entropy Generation Rate(EGR)is used to analyze the irreversible loss.It turns out that the SWWVI can reduce the flow field loss.There are several weak shock waves in the NGV flow field,which can increase the irreversible loss.This work offers flow mechanisms analysis and presents the EGR distribution in SWBLI and SWWVI areas in a transonic turbine blade.展开更多
Three-dimensional curved shock wave/boundary layer interaction with streamwise and spanwise curvatures widely exists in practical aerodynamic design.To explore the effects of composite shock curvatures on boundary lay...Three-dimensional curved shock wave/boundary layer interaction with streamwise and spanwise curvatures widely exists in practical aerodynamic design.To explore the effects of composite shock curvatures on boundary layer separation,a canonical model with a cone placed above plate was utilized as a reference.Configurations of straight,convex,and concave conical shock waves inducing the curved conical shock wave/boundary layer interactions were studied,using CFD based on Reynolds-averaged numerical simulation method.The flow structure and separation region of each case were discussed quantitively on the symmetry plane,flat plate,and plane perpendicular to flow direction,respectively.The focus of the analysis was on the characteristic patterns of separation scale variation in the streamwise and spanwise directions,which were observed to consistently change with respect to both directions with alterations in the incident shock wave shape.A simplified control volume model was established to qualitatively discuss the influence source of curved shock waves on separation scales,based on mass conservation equations.The results suggest that the curved shock wave has a holistic effect on separation,which is not solely dependent on the shock foot strength.展开更多
Cowl-induced incident Shock Wave/Boundary Layer Interactions (SWBLI) under the influence of gradual expansion waves are frequently observed in supersonic inlets. However, the analysis and prediction of interaction len...Cowl-induced incident Shock Wave/Boundary Layer Interactions (SWBLI) under the influence of gradual expansion waves are frequently observed in supersonic inlets. However, the analysis and prediction of interaction lengths have not been sufficiently investigated. First, this study presents a theoretical scaling analysis and validates it through wind tunnel experiments. It conducts detailed control volume analysis of mass conservation, considering the differences between inviscid and viscous cases. Then, three models for analysing interaction length under gradual expansion waves are derived. Related experiments using schlieren photography are conducted to validate the models in a Mach 2.73 flow. The interaction scales are captured at various relative distances between the shock impingement location and the expansion regions with wedge angles ranging from 12° to 15° and expansion angles of 9°, 12°, and 15°. Three trend lines are plotted based on different expansion angles to depict the relationship between normalised interaction length and normalised interaction strength metric. In addition, the relationship between the coefficients of the trend line and the expansion angles is introduced to predict the interaction length influenced by gradual expansion waves. Finally, the estimation of normalised interaction length is derived for various coefficients within a unified form.展开更多
A Discrete Boltzmann Method(DBM)with a Maxwell-type boundary condition is constructed to investigate the influence of rarefaction on laminar Shock Wave/Boundary Layer Interaction(SWBLI).Due to the complexity of compre...A Discrete Boltzmann Method(DBM)with a Maxwell-type boundary condition is constructed to investigate the influence of rarefaction on laminar Shock Wave/Boundary Layer Interaction(SWBLI).Due to the complexity of compressible flow,a Knudsen number vector Kn,whose components include the local Knudsen numbers such as Kn_(ρ)and Kn_(U),is introduced to characterize the local structures,where Kn_(ρ)and Kn_(U)are Knudsen numbers defined in terms of the density and velocity interfaces,respectively.Since first focusing on the steady state of SWBLI,the DBM considers up to the second-order Kn_(ρ)(rarefaction/non-equilibrium)effects.The model is validated using Mach number 2 SWBLI and the necessity of using DBM with sufficient physical accuracy is confirmed by the shock collision problem.Key findings include the following:the leading-edge shock wave increases the local density Knudsen number Kn_(ρ)and eventually leads to the failure of linear constitutive relations in the Navier-Stokes(N-S)model and surely also in the lower-order DBM;the non-equilibrium effect differences in regions behind the leading-edge shock wave are primarily correlated with Kn_(ρ),while in the separation region are primarily correlated with Kn_(U);the non-equilibrium quantities D_(2)and D_(4,2),as well as the viscous entropy production rate S_(NOMF)can be used to identify the separation zone.The findings clarify various effects and main mechanisms in different regions associated with SWBLI,which are concealed in N-S model.展开更多
The phenomenon of shock/shock interaction(SSI)is widely observed in high-speed flow,and the double wedge SSI represents one of the typical problems encountered.The control effect of single-pulse plasma synthetic jet(P...The phenomenon of shock/shock interaction(SSI)is widely observed in high-speed flow,and the double wedge SSI represents one of the typical problems encountered.The control effect of single-pulse plasma synthetic jet(PSJ)on double wedge type-Ⅵand type-ⅤSSI was investigated experimentally and numerically,and the influence of discharge energy was also explored.The findings indicate that the interaction between PSJ and the high-speed freestream results in the formation of a plasma layer and a jet shock,which collectively governs the control of SSI.The control mechanism of single-pulse PSJ on SSI lies in its capacity to attenuate both shock and SSI.For type-ⅥSSI,the original second-wedge oblique shock is eliminated under the control of PSJ,resulting in a new type-ⅥSSI formed by the jet shock and the first-wedge oblique shock.For type-ⅤSSI,the presence of PSJ effectively mitigates the intensity of Mach stem,supersonic jet,and reflected shocks,thereby facilitating its transition into type-ⅥSSI.The numerical results indicate that the peak pressure can be reduced by approximately 32.26%at maximum.Furthermore,the development of PSJ also extends in the Z direction.The pressure decreases in the area affected by both PSJ and jet shock due to the attenuation of the SSI zone.With increasing discharge energy,the control effect of PSJ on SSI is gradually enhanced.展开更多
A two-dimensional Reynolds-averaged Navier-Stokes solver is applied to analyze the aerodynamic behavior of the Shock/Boundary-Layer interaction of rocket with a boosted The K-ε turbulence model and a finite volume m...A two-dimensional Reynolds-averaged Navier-Stokes solver is applied to analyze the aerodynamic behavior of the Shock/Boundary-Layer interaction of rocket with a boosted The K-ε turbulence model and a finite volume method in a unstructured body-fitted curvilinear coordinates have been used. The results indicate that the separation and the reattachment occur in the Boundary-Layer of the main rocket because of the shock interaction. The shape of the booster nose effects the flow field obviously. In the case of the hemisphere booster nose the pressure has complicate distributions and the separation is very clear. The distance between the booster and main rocket has the evident effect on the flow field. If the distance is smaller the pressure coefficient is bigger the separation zone even the separation bubble occurs.展开更多
This paper examines the Shock/Shock Interactions（SSI）between the body and wing of aircraft in supersonic flows.The body is simplified to a flat wedge and the wing is assumed to be a sharp wing.The theoretical spatia...This paper examines the Shock/Shock Interactions（SSI）between the body and wing of aircraft in supersonic flows.The body is simplified to a flat wedge and the wing is assumed to be a sharp wing.The theoretical spatial dimension reduction method,which transforms the 3D problem into a 2D one,is used to analyze the SSI between the body and wing.The temperature and pressure behind the Mach stem induced by the wing and body are obtained,and the wave configurations in the corner are determined.Numerical validations are conducted by solving the inviscid Euler equations in 3D with a Non-oscillatory and Non-free-parameters Dissipative（NND）finite difference scheme.Good agreements between the theoretical and numerical results are obtained.Additionally,the effects of the wedge angle and sweep angle on wave configurations and flow field are considered numerically and theoretically.The influences of wedge angle are significant,whereas the effects of sweep angle on wave configurations are negligible.This paper provides useful information for the design and thermal protection of aircraft in supersonic and hypersonic flows.展开更多
An experimental study was conducted on shock wave turbulent boundary layer interactions caused by a blunt swept fin-plate configuration at Mach numbers of 5.0, 7.8, 9.9 for a Reynolds number range of (1.0.similar to 4...An experimental study was conducted on shock wave turbulent boundary layer interactions caused by a blunt swept fin-plate configuration at Mach numbers of 5.0, 7.8, 9.9 for a Reynolds number range of (1.0.similar to 4.7) x 10(7)/m. Detailed heat transfer and pressure distributions were measured at fin deflection angles of up to 30 degrees for a sweepback angle of 67.6 degrees. Surface oil flow patterns and liquid crystal thermograms as well as schlieren pictures of fin shock shape were taken. The study shows that the flow was separated at deflection of 10 degrees and secondary separation were detected at deflection of theta greater than or equal to 20 degrees. The heat transfer and pressure distributions on flat plate showed an extensive plateau region followed by a distinct dip and local peak close to the fin foot. Measurements of the plateau pressure and heat transfer were in good agreement with existing prediction methods, but pressure and heating peak measurements at M greater than or equal to 6 were significantly lower than predicted by the simple prediction techniques at lower Mach numbers.展开更多
The reason for the asymmetry phenomenon of shock/boundary layer interactions(SBLI)in a completely symmetric nozzle with symmetric flow conditions is still an open question.A model for the asymmetry of nozzle flows was...The reason for the asymmetry phenomenon of shock/boundary layer interactions(SBLI)in a completely symmetric nozzle with symmetric flow conditions is still an open question.A model for the asymmetry of nozzle flows was proposed based on the properties of fluid entrainment in the mixing layer and momentum conservation.The asymmetry model is deduced based on the nozzle flow with restricted shock separation,and is still applicable for free shock separation.Flow deflection angle at nozzle exit is deduced from this model.Steady numerical simulations are conducted to model the asymmetry of the SBLIs in a planar convergent-divergent nozzle tested by previous researchers.The obtained values of deflection angle based on the numerical results of forced symmetric nozzle flows can judge the asymmetry of flows in a nozzle at some operations.It shows that the entrainment of shear layer on the separation induced by SBLTs is one of the reasons for the asymmetry in the confined SBLIs.展开更多
Observations are presented from experiments and calculations where a laminar spherical CH4/air flame is perturbed successively by incident and reflected shock waves. The experiments are performed in a standard shock t...Observations are presented from experiments and calculations where a laminar spherical CH4/air flame is perturbed successively by incident and reflected shock waves. The experiments are performed in a standard shock tube arrangement, in which a high-speed shadowgraph imaging system is used to record evolutions of the flame. Numerical simulations are conducted by using second-order wave propagation algorithms, based on two-dimensional axisymmetric Navier-Stokes equations with detailed chemical reactions. Qualitative agreements are obtained between the experimental and numerical results. Under actions of incident shock waves, Richtmyer-Meshkov instability responsible for the flame deformation is induced in the flame, and the distoned flame takes a barrel shape. Then, under subsequent actions of the shock wave reflected from a planar wall, the flame takes an inclined non-symmetrical kidney shape in a symmetric cross section, which means a mushroom-like shape of the flame comes finally into being. The vorticity direction in the ring cap has been altered by the reflected shock＇s action, which makes the head of the mushroom-like flame extend quickly to the side wall.展开更多
Shock formation due to flow compressibility and its interaction with boundary layers has adverse effects on aerodynamic characteristics, such as drag increase and flow separation. The objective of this paper is to app...Shock formation due to flow compressibility and its interaction with boundary layers has adverse effects on aerodynamic characteristics, such as drag increase and flow separation. The objective of this paper is to appraise the practicability of weakening shock waves and, hence, reducing the wave drag in transonic flight regime using a two-dimensional jagged wall and thereby to gain an appropriate jagged wall shape for future empirical study. Different shapes of the jagged wall, including rectangular, circular, and triangular shapes, were employed. The numerical method was validated by experimental and numerical studies involving transonic flow over the NACA0012 airfoil, and the results presented here closely match previous experimental and numerical results. The impact of parameters, including shape and the length-to-spacing ratio of a jagged wall, was studied on aerodynamic forces and flow field. The results revealed that applying a jagged wall method on the upper surface of an airfoil changes the shock structure significantly and disintegrates it, which in turn leads to a decrease in wave drag. It was also found that the maximum drag coefficient decrease of around 17 % occurs with a triangular shape, while the maximum increase in aerodynamic efficiency（lift-to-drag ratio）of around 10 % happens with a rectangular shape at an angle of attack of 2.26？.展开更多
An experimental study and a numerical simulation were conducted to investigate the mechanical and thermodynamic processes involved in the interaction between shock waves and low density foam. The experiment was done i...An experimental study and a numerical simulation were conducted to investigate the mechanical and thermodynamic processes involved in the interaction between shock waves and low density foam. The experiment was done in a stainless shock tube (80 mm in inner diameter, 10 mm in wall thickness and 5 360 mm in length). The velocities of the incident and reflected compression waves in the foam were measured by using piezo-ceramic pressure sensors. The end-wall peak pressure behind the reflected wave in the foam was measured by using a crystal piezoelectric sensor. It is suggested that the high end-wall pressure may be caused by a rapid contact between the foam and the end-wall surface. Both open-cell and closed-cell foams with different length and density were tested. Through comparing the numerical and experimental end-wall pressure, the permeability coefficients α and β are quantitatively determined.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
The reflection and diffraction of a planar shock wave around a circular cylinder are a typical problem of the complex nonlinear shock wave phenomena in literature.It has long been studied experimentally,analytically a...The reflection and diffraction of a planar shock wave around a circular cylinder are a typical problem of the complex nonlinear shock wave phenomena in literature.It has long been studied experimentally,analytically as well as numerically.Takayama in 1987 obtained clear experimental pictures ofisopycnics in shock tube under the condi- tion that the impinging shock wave propagates as far as 3 diameters away from the cylinder.To know more complete- ly the whole unsteady process,it is desirable to get experimental results in a region which is more than 10 diameters away from the cylinder.This is what has been done in this paper by using the pulsed laser holographic interferometry for several shock Mach numbers of the impinging shock. Results for several moments are shown,giving more know- ledge about the whole unsteady flow field.This is useful for a reliable and complete understanding of the changing force acting on the cylinder,and provides interesting data to check the performance of many recently developed high resolution numerical methods for unsteady shock wave calculation.展开更多
The properties of Mach stems in hypersonic corner flow induced by Mach interaction over 3D intersecting wedges were studied theoretically and numerically.A new method called ＂spatial dimension reduction＂ was used to...The properties of Mach stems in hypersonic corner flow induced by Mach interaction over 3D intersecting wedges were studied theoretically and numerically.A new method called ＂spatial dimension reduction＂ was used to analyze theoretically the location and Mach number behind Mach stems. By using this approach, the problem of 3D steady shock/shock interaction over 3D intersecting wedges was transformed into a 2D moving one on cross sections, which can be solved by shock-polar theory and shock dynamics theory. The properties of Mach interaction over 3D intersecting wedges can be analyzed with the new method,including pressure, temperature, density in the vicinity of triple points, location, and Mach number behind Mach stems.Theoretical results were compared with numerical results,and good agreement was obtained. Also, the influence of Mach number and wedge angle on the properties of a 3D Mach stem was studied.展开更多
Based on the working principle and the damping characteristic of hydraulic shock absorber, a fluid structure interaction method was presented, which was used to analyze the microcosmic and high-frequency processing me...Based on the working principle and the damping characteristic of hydraulic shock absorber, a fluid structure interaction method was presented, which was used to analyze the microcosmic and high-frequency processing mechanism of fluid structure interaction between circulation valve and liquid of hydraulic shock absorber. The fluid mesh distortion was controlled by the CEL language, and the fluid struc^tre interaction mathematical model was established. The finite element model was established by ANSYS CFX software and was analyzed by dynamic mesh technique. The local sensitive computational area was meshed by prismatic grid, which could reduce the negative volume problem during the simulation. The circulation valve and liquid of hydraulic shock absorber were simulated and analyzed under the condition of sinusoidal inlet velocity loads. Flow characteristic and dynamics characteristic were obtained. The pressure distribution and the displacement of circulation value were obtained, and the acceleration curve of circulation valve was simulated and analyzed. The conformity of the final simulation results with the experimental datum indicates that this method is accurate and reliable to analyze the dynamics characteristic between circulation valve and liquid of hydraulic shock absorber, which can provide a theoretical foundation for optimizing hydraulic shock absorber in the future.展开更多
The interaction length induced by Shock Wave/Turbulent Boundary-Layer Interactions(SWTBLIs)in the hypersonic flow was investigated using a scaling analysis,in which the interaction length normalized by the displacemen...The interaction length induced by Shock Wave/Turbulent Boundary-Layer Interactions(SWTBLIs)in the hypersonic flow was investigated using a scaling analysis,in which the interaction length normalized by the displacement thickness of boundary layer was correlated with a corrected non-dimensional separation criterion across the interaction after accounting for the wall temperature effects.A large number of hypersonic SWTBLIs were compiled to examine the scaling analysis over a wide range of Mach numbers,Reynolds numbers,and wall temperatures.The results indicate that the hypersonic SWTBLIs with low Reynolds numbers collapse on the supersonic SWTBLIs,while the hypersonic cases with high Reynolds numbers show a more rapid growth of the interaction length than that with low Reynolds numbers.Thus,two scaling relationships are identified according to different Reynolds numbers for the hypersonic SWTBLIs.The scaling analysis provides valuable guidelines for engineering prediction of the interaction length,and thus,enriches the knowledge of hypersonic SWTBLIs.展开更多
基金supported by the National Natural Science Foundation of China(No.12372233)the Innovation Foundation for Doctor Dissertation of Northwestern Polytechnical University,China(No.25GH01020005)the“111 Project”of China(No.B17037)。
文摘As a multidisciplinary phenomenon,panel aeroelasticity in shock-dominated flow is featured by two primary interactions:Fluid-Structure Interactions(FSIs)and Shock-Boundary Layer Interactions(SBLIs).The former raises structural concerns,and the latter is of aerodynamic interest.Thus,panel aeroelasticity in shock-dominated flow represents a vital topic for the development and optimization of supersonic vehicles and propulsion systems.This review systematically summarizes recent advances in the methodologies applied to capture structural and fluid dynamics,including theoretical models,numerical simulations,and wind tunnel experiments.The application of data-driven modal decomposition,an advanced technique to extract physically crucial features,on the topic is introduced.From the perspective of FSIs,the distinctive aeroelastic behaviors in shock-dominated flow,including hysteresis phenomena and nonlinear responses,are highlighted.From the perspective of SBLIs,the modifications in their spatial and temporal characteristics imposed by the aeroelastic responses are emphasized.Motivated by the interaction between the shock waves and structural response,different strategies have been proposed to implement aeroelastic suppression and shock control,which have the potential to enhance structural safety and aerodynamic performance in the next generation of high-speed flight vehicles.
基金funded by the Research Fund of National Key Laboratory of Aerospace Physics in Fluids,grant number 2024-APF-KFZD-01Guangdong Basic and Applied Basic Research Foundation,grant number 2025A1515012081+1 种基金National Natural Science Foundation of China,grant number 12002193Shandong Provincial Natural Science Foundation,China,grant number ZR2019QA018.
文摘For hypersonic air-breathing vehicles,the V-shaped leading edges(VSLEs)of supersonic combustion ramjet(scramjet)inlets experience complex shock interactions and intense aerodynamic loads.This paper provides a comprehensive review of flow characteristics at the crotch of VSLEs,with particular focus on the transition of shock interaction types and the variation of wall heat flux under different freestream Mach numbers and geometric configurations.The mechanisms governing shock transition,unsteady oscillations,hysteresis,and three-dimensional effects in VSLE flows are first examined.Subsequently,thermal protection strategies aimed at mitigating extreme heating loads are reviewed,emphasizing their relevance to practical engineering applications.Special attention is given to recent studies addressing thermochemical nonequilibrium effects on VSLE shock interactions,and the limitations of current research are critically assessed.Finally,perspectives for future investigations into hypersonic VSLE shock interactions are outlined,highlighting opportunities for advancing design and thermal management strategies.
基金supported by the National Key R&D Program of China(Grant No.2019YFA0405300)the National Natural Science Foundation of China(Grant No.11972368)the Natural Science Foundation of Hunan Province(Grant No.2021JJ10045).
文摘The stability of supersonic inlets faces challenges due to various changes in flight conditions,and flow control methods that address shock wave/boundary layer interactions under only one set of conditions cannot meet developmental requirements.This paper proposes an adaptive bump control scheme and employs dynamic mesh technology for numerical simulation to investigate the unsteady control effects of adaptive bumps.The obtained results indicate that the use of moving bumps to control shock wave/boundary layer interactions is feasible.The adaptive control effects of five different bump speeds are evaluated.Within the range of bump speeds studied,the analysis of the flow field structure reveals the patterns of change in the separation zone area during the control process,as well as the relationship between the bump motion speed and the control effect on the separation zone.It is concluded that the moving bump endows the boundary layer with additional energy.
基金supported by the National Natural Science Foundation of China(No.51876098).
文摘Accurate predictions of Shock Waves and Boundary Layer Interaction(SWBLI)and strong Shock Waves and Wake Vortices Interaction(SWWVI)in a highly-loaded turbine propose challenges to the currently widely used Reynolds-Averaged Navier-Stokes(RANS)model.In this work,the SWBLI and the SWWVI in a highly-loaded Nozzle Guide Vane(NGV)are studied using a hybrid RANS/LES strategy.The Turbulence Kinetic Energy(TKE)budget and the Proper Orthogonal Decomposition(POD)method are used to analyze flow mechanisms.Results show that this hybrid RANS/LES method can obtain detailed flow structures for flow mechanisms analysis.Strong shock waves induce boundary layer separation,while the presence of a separation bubble can in turn lead to a Mach reflection phenomenon.The shock waves cause trailing-edge vortices to break clearly,and the wakes,in turn,can change the shocks intensity and direction.Furthermore,the Entropy Generation Rate(EGR)is used to analyze the irreversible loss.It turns out that the SWWVI can reduce the flow field loss.There are several weak shock waves in the NGV flow field,which can increase the irreversible loss.This work offers flow mechanisms analysis and presents the EGR distribution in SWBLI and SWWVI areas in a transonic turbine blade.
基金the support of the National Natural Science Foundation of China(Nos.12372295,U21B6003,U20A2069,12302389 and 123B2037)。
文摘Three-dimensional curved shock wave/boundary layer interaction with streamwise and spanwise curvatures widely exists in practical aerodynamic design.To explore the effects of composite shock curvatures on boundary layer separation,a canonical model with a cone placed above plate was utilized as a reference.Configurations of straight,convex,and concave conical shock waves inducing the curved conical shock wave/boundary layer interactions were studied,using CFD based on Reynolds-averaged numerical simulation method.The flow structure and separation region of each case were discussed quantitively on the symmetry plane,flat plate,and plane perpendicular to flow direction,respectively.The focus of the analysis was on the characteristic patterns of separation scale variation in the streamwise and spanwise directions,which were observed to consistently change with respect to both directions with alterations in the incident shock wave shape.A simplified control volume model was established to qualitatively discuss the influence source of curved shock waves on separation scales,based on mass conservation equations.The results suggest that the curved shock wave has a holistic effect on separation,which is not solely dependent on the shock foot strength.
基金co-supported by the National Natural Science Foundation of China (No. 12172175)the National Science and Technology Major Project, China (No. J2019-II0014-0035)the Science Center for Gas Turbine Project, China (Nos. P2022-C-II-002-001, P2022-A-II-002-001)
文摘Cowl-induced incident Shock Wave/Boundary Layer Interactions (SWBLI) under the influence of gradual expansion waves are frequently observed in supersonic inlets. However, the analysis and prediction of interaction lengths have not been sufficiently investigated. First, this study presents a theoretical scaling analysis and validates it through wind tunnel experiments. It conducts detailed control volume analysis of mass conservation, considering the differences between inviscid and viscous cases. Then, three models for analysing interaction length under gradual expansion waves are derived. Related experiments using schlieren photography are conducted to validate the models in a Mach 2.73 flow. The interaction scales are captured at various relative distances between the shock impingement location and the expansion regions with wedge angles ranging from 12° to 15° and expansion angles of 9°, 12°, and 15°. Three trend lines are plotted based on different expansion angles to depict the relationship between normalised interaction length and normalised interaction strength metric. In addition, the relationship between the coefficients of the trend line and the expansion angles is introduced to predict the interaction length influenced by gradual expansion waves. Finally, the estimation of normalised interaction length is derived for various coefficients within a unified form.
基金support from the National Key R&D Program of China(No.2020YFC2201100)the Foundation of National Key Laboratory of Shock Wave and Detonation Physics,China(No.JCKYS2023212003)+1 种基金the National Natural Science Foundation of China(No.12172061)the Opening Project of State Key Laboratory of Explosion Science and Safety Protection(Beijing Institute of Technology)(No.KFJJ25-02M).
文摘A Discrete Boltzmann Method(DBM)with a Maxwell-type boundary condition is constructed to investigate the influence of rarefaction on laminar Shock Wave/Boundary Layer Interaction(SWBLI).Due to the complexity of compressible flow,a Knudsen number vector Kn,whose components include the local Knudsen numbers such as Kn_(ρ)and Kn_(U),is introduced to characterize the local structures,where Kn_(ρ)and Kn_(U)are Knudsen numbers defined in terms of the density and velocity interfaces,respectively.Since first focusing on the steady state of SWBLI,the DBM considers up to the second-order Kn_(ρ)(rarefaction/non-equilibrium)effects.The model is validated using Mach number 2 SWBLI and the necessity of using DBM with sufficient physical accuracy is confirmed by the shock collision problem.Key findings include the following:the leading-edge shock wave increases the local density Knudsen number Kn_(ρ)and eventually leads to the failure of linear constitutive relations in the Navier-Stokes(N-S)model and surely also in the lower-order DBM;the non-equilibrium effect differences in regions behind the leading-edge shock wave are primarily correlated with Kn_(ρ),while in the separation region are primarily correlated with Kn_(U);the non-equilibrium quantities D_(2)and D_(4,2),as well as the viscous entropy production rate S_(NOMF)can be used to identify the separation zone.The findings clarify various effects and main mechanisms in different regions associated with SWBLI,which are concealed in N-S model.
基金supported by the Independent Innovation Science Fund of National University of Defense Technology(No.24-ZZCX-BC-05)National Natural Science Foundation of China(Nos.92271110 and 12202488)+2 种基金the Major National Science and Technology Project(No.J2019-Ⅲ0010-0054)the National Postdoctoral Researcher Program of China(No.GZB20230985)the Natural Science Program of National University of Defense Technology(No.ZK22-30)。
文摘The phenomenon of shock/shock interaction(SSI)is widely observed in high-speed flow,and the double wedge SSI represents one of the typical problems encountered.The control effect of single-pulse plasma synthetic jet(PSJ)on double wedge type-Ⅵand type-ⅤSSI was investigated experimentally and numerically,and the influence of discharge energy was also explored.The findings indicate that the interaction between PSJ and the high-speed freestream results in the formation of a plasma layer and a jet shock,which collectively governs the control of SSI.The control mechanism of single-pulse PSJ on SSI lies in its capacity to attenuate both shock and SSI.For type-ⅥSSI,the original second-wedge oblique shock is eliminated under the control of PSJ,resulting in a new type-ⅥSSI formed by the jet shock and the first-wedge oblique shock.For type-ⅤSSI,the presence of PSJ effectively mitigates the intensity of Mach stem,supersonic jet,and reflected shocks,thereby facilitating its transition into type-ⅥSSI.The numerical results indicate that the peak pressure can be reduced by approximately 32.26%at maximum.Furthermore,the development of PSJ also extends in the Z direction.The pressure decreases in the area affected by both PSJ and jet shock due to the attenuation of the SSI zone.With increasing discharge energy,the control effect of PSJ on SSI is gradually enhanced.
文摘A two-dimensional Reynolds-averaged Navier-Stokes solver is applied to analyze the aerodynamic behavior of the Shock/Boundary-Layer interaction of rocket with a boosted The K-ε turbulence model and a finite volume method in a unstructured body-fitted curvilinear coordinates have been used. The results indicate that the separation and the reattachment occur in the Boundary-Layer of the main rocket because of the shock interaction. The shape of the booster nose effects the flow field obviously. In the case of the hemisphere booster nose the pressure has complicate distributions and the separation is very clear. The distance between the booster and main rocket has the evident effect on the flow field. If the distance is smaller the pressure coefficient is bigger the separation zone even the separation bubble occurs.
基金supported by the Fundamental Research Funds for the Central Universities of China (No. 31020170QD087)
文摘This paper examines the Shock/Shock Interactions（SSI）between the body and wing of aircraft in supersonic flows.The body is simplified to a flat wedge and the wing is assumed to be a sharp wing.The theoretical spatial dimension reduction method,which transforms the 3D problem into a 2D one,is used to analyze the SSI between the body and wing.The temperature and pressure behind the Mach stem induced by the wing and body are obtained,and the wave configurations in the corner are determined.Numerical validations are conducted by solving the inviscid Euler equations in 3D with a Non-oscillatory and Non-free-parameters Dissipative（NND）finite difference scheme.Good agreements between the theoretical and numerical results are obtained.Additionally,the effects of the wedge angle and sweep angle on wave configurations and flow field are considered numerically and theoretically.The influences of wedge angle are significant,whereas the effects of sweep angle on wave configurations are negligible.This paper provides useful information for the design and thermal protection of aircraft in supersonic and hypersonic flows.
基金The project supported by China Academy of Launch Vehicle Technology
文摘An experimental study was conducted on shock wave turbulent boundary layer interactions caused by a blunt swept fin-plate configuration at Mach numbers of 5.0, 7.8, 9.9 for a Reynolds number range of (1.0.similar to 4.7) x 10(7)/m. Detailed heat transfer and pressure distributions were measured at fin deflection angles of up to 30 degrees for a sweepback angle of 67.6 degrees. Surface oil flow patterns and liquid crystal thermograms as well as schlieren pictures of fin shock shape were taken. The study shows that the flow was separated at deflection of 10 degrees and secondary separation were detected at deflection of theta greater than or equal to 20 degrees. The heat transfer and pressure distributions on flat plate showed an extensive plateau region followed by a distinct dip and local peak close to the fin foot. Measurements of the plateau pressure and heat transfer were in good agreement with existing prediction methods, but pressure and heating peak measurements at M greater than or equal to 6 were significantly lower than predicted by the simple prediction techniques at lower Mach numbers.
基金supported by the National Natural Science Foundations of China(Nos.51476076,51776096)
文摘The reason for the asymmetry phenomenon of shock/boundary layer interactions(SBLI)in a completely symmetric nozzle with symmetric flow conditions is still an open question.A model for the asymmetry of nozzle flows was proposed based on the properties of fluid entrainment in the mixing layer and momentum conservation.The asymmetry model is deduced based on the nozzle flow with restricted shock separation,and is still applicable for free shock separation.Flow deflection angle at nozzle exit is deduced from this model.Steady numerical simulations are conducted to model the asymmetry of the SBLIs in a planar convergent-divergent nozzle tested by previous researchers.The obtained values of deflection angle based on the numerical results of forced symmetric nozzle flows can judge the asymmetry of flows in a nozzle at some operations.It shows that the entrainment of shear layer on the separation induced by SBLTs is one of the reasons for the asymmetry in the confined SBLIs.
基金supported by the National Natural Science Foundation of China (10472047)the Open Fund of State Key Laboratory of Explosion Science Technology, Beijing University of Science and Technology (KFJJ06-3)
文摘Observations are presented from experiments and calculations where a laminar spherical CH4/air flame is perturbed successively by incident and reflected shock waves. The experiments are performed in a standard shock tube arrangement, in which a high-speed shadowgraph imaging system is used to record evolutions of the flame. Numerical simulations are conducted by using second-order wave propagation algorithms, based on two-dimensional axisymmetric Navier-Stokes equations with detailed chemical reactions. Qualitative agreements are obtained between the experimental and numerical results. Under actions of incident shock waves, Richtmyer-Meshkov instability responsible for the flame deformation is induced in the flame, and the distoned flame takes a barrel shape. Then, under subsequent actions of the shock wave reflected from a planar wall, the flame takes an inclined non-symmetrical kidney shape in a symmetric cross section, which means a mushroom-like shape of the flame comes finally into being. The vorticity direction in the ring cap has been altered by the reflected shock＇s action, which makes the head of the mushroom-like flame extend quickly to the side wall.
文摘Shock formation due to flow compressibility and its interaction with boundary layers has adverse effects on aerodynamic characteristics, such as drag increase and flow separation. The objective of this paper is to appraise the practicability of weakening shock waves and, hence, reducing the wave drag in transonic flight regime using a two-dimensional jagged wall and thereby to gain an appropriate jagged wall shape for future empirical study. Different shapes of the jagged wall, including rectangular, circular, and triangular shapes, were employed. The numerical method was validated by experimental and numerical studies involving transonic flow over the NACA0012 airfoil, and the results presented here closely match previous experimental and numerical results. The impact of parameters, including shape and the length-to-spacing ratio of a jagged wall, was studied on aerodynamic forces and flow field. The results revealed that applying a jagged wall method on the upper surface of an airfoil changes the shock structure significantly and disintegrates it, which in turn leads to a decrease in wave drag. It was also found that the maximum drag coefficient decrease of around 17 % occurs with a triangular shape, while the maximum increase in aerodynamic efficiency（lift-to-drag ratio）of around 10 % happens with a rectangular shape at an angle of attack of 2.26？.
文摘An experimental study and a numerical simulation were conducted to investigate the mechanical and thermodynamic processes involved in the interaction between shock waves and low density foam. The experiment was done in a stainless shock tube (80 mm in inner diameter, 10 mm in wall thickness and 5 360 mm in length). The velocities of the incident and reflected compression waves in the foam were measured by using piezo-ceramic pressure sensors. The end-wall peak pressure behind the reflected wave in the foam was measured by using a crystal piezoelectric sensor. It is suggested that the high end-wall pressure may be caused by a rapid contact between the foam and the end-wall surface. Both open-cell and closed-cell foams with different length and density were tested. Through comparing the numerical and experimental end-wall pressure, the permeability coefficients α and β are quantitatively determined.
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.
基金The project suported partially by National Natural Science Foundation of China
文摘The reflection and diffraction of a planar shock wave around a circular cylinder are a typical problem of the complex nonlinear shock wave phenomena in literature.It has long been studied experimentally,analytically as well as numerically.Takayama in 1987 obtained clear experimental pictures ofisopycnics in shock tube under the condi- tion that the impinging shock wave propagates as far as 3 diameters away from the cylinder.To know more complete- ly the whole unsteady process,it is desirable to get experimental results in a region which is more than 10 diameters away from the cylinder.This is what has been done in this paper by using the pulsed laser holographic interferometry for several shock Mach numbers of the impinging shock. Results for several moments are shown,giving more know- ledge about the whole unsteady flow field.This is useful for a reliable and complete understanding of the changing force acting on the cylinder,and provides interesting data to check the performance of many recently developed high resolution numerical methods for unsteady shock wave calculation.
基金supported by the National Natural Science Foundation of China (Grants 11372333, 90916028)
文摘The properties of Mach stems in hypersonic corner flow induced by Mach interaction over 3D intersecting wedges were studied theoretically and numerically.A new method called ＂spatial dimension reduction＂ was used to analyze theoretically the location and Mach number behind Mach stems. By using this approach, the problem of 3D steady shock/shock interaction over 3D intersecting wedges was transformed into a 2D moving one on cross sections, which can be solved by shock-polar theory and shock dynamics theory. The properties of Mach interaction over 3D intersecting wedges can be analyzed with the new method,including pressure, temperature, density in the vicinity of triple points, location, and Mach number behind Mach stems.Theoretical results were compared with numerical results,and good agreement was obtained. Also, the influence of Mach number and wedge angle on the properties of a 3D Mach stem was studied.
基金Project(51275542) supported by the National Natural Science Foundation of Chinaproject(CDJXS12110010) supported by the Fundamental Research Funds for the Central Universities of China
文摘Based on the working principle and the damping characteristic of hydraulic shock absorber, a fluid structure interaction method was presented, which was used to analyze the microcosmic and high-frequency processing mechanism of fluid structure interaction between circulation valve and liquid of hydraulic shock absorber. The fluid mesh distortion was controlled by the CEL language, and the fluid struc^tre interaction mathematical model was established. The finite element model was established by ANSYS CFX software and was analyzed by dynamic mesh technique. The local sensitive computational area was meshed by prismatic grid, which could reduce the negative volume problem during the simulation. The circulation valve and liquid of hydraulic shock absorber were simulated and analyzed under the condition of sinusoidal inlet velocity loads. Flow characteristic and dynamics characteristic were obtained. The pressure distribution and the displacement of circulation value were obtained, and the acceleration curve of circulation valve was simulated and analyzed. The conformity of the final simulation results with the experimental datum indicates that this method is accurate and reliable to analyze the dynamics characteristic between circulation valve and liquid of hydraulic shock absorber, which can provide a theoretical foundation for optimizing hydraulic shock absorber in the future.
基金supported by the National Natural Science Foundation of China(Nos.11772325 and 11621202)。
文摘The interaction length induced by Shock Wave/Turbulent Boundary-Layer Interactions(SWTBLIs)in the hypersonic flow was investigated using a scaling analysis,in which the interaction length normalized by the displacement thickness of boundary layer was correlated with a corrected non-dimensional separation criterion across the interaction after accounting for the wall temperature effects.A large number of hypersonic SWTBLIs were compiled to examine the scaling analysis over a wide range of Mach numbers,Reynolds numbers,and wall temperatures.The results indicate that the hypersonic SWTBLIs with low Reynolds numbers collapse on the supersonic SWTBLIs,while the hypersonic cases with high Reynolds numbers show a more rapid growth of the interaction length than that with low Reynolds numbers.Thus,two scaling relationships are identified according to different Reynolds numbers for the hypersonic SWTBLIs.The scaling analysis provides valuable guidelines for engineering prediction of the interaction length,and thus,enriches the knowledge of hypersonic SWTBLIs.
