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
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
The interactions of oblique/bow shock waves are the key flow phenomena restricting the design and aerothermodynamic performance of high-speed vehicles.Type Ⅲ and Type Ⅳ Shock/Shock Interactions(SSIs)have been extens...The interactions of oblique/bow shock waves are the key flow phenomena restricting the design and aerothermodynamic performance of high-speed vehicles.Type Ⅲ and Type Ⅳ Shock/Shock Interactions(SSIs)have been extensively investigated,as such interactions can induce abnormal aerodynamic heating problems in hypersonic flows of vehicles.The transition process between these two distinct types of shock/shock interactions remains unclear.In the present study,a subclass of shock/shock interaction configuration is revealed and defined as Type Ⅲa.Type Ⅲa interaction can induce much more severe aerodynamic heating than a Type Ⅳ interaction which was ever reported to be the most serious in literature.The intense aerodynamic heating observed in this configuration highlights a new design point for the thermal protection system of hypersonic vehicles.A secondary Mach interaction between shock waves in the supersonic flow path of a Type Ⅲ configuration is demonstrated to be the primary mechanism for such a subclass of shock/shock interaction configuration.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
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 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
A shock interaction problem is solved with finite difference methods for a hypersonic fiow of air with chemical reactions. If a body has two concave cornerst a secondary shock is formed in the shock layer and it meets...A shock interaction problem is solved with finite difference methods for a hypersonic fiow of air with chemical reactions. If a body has two concave cornerst a secondary shock is formed in the shock layer and it meets the main shock later.As the two shocks meet, the flow becomes singular at the interaction point, and a new main shock, a contact discontinuity and an expansion wave appear as a result of interaction between the two shocks. Therefore, the problem is very complicated.Using proper combinations of implicit and explicit finite difference schemes according to the property of the equations and the boundary conditions, we compute the fiow behind the interaction point successfully.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
This paper explores theoretical solutions to the three-dimensional （3D） shock/shock interaction induced by 3D asymmetrical intersecting compression wedges in supersonic inviscid flows. For Mach interactions, an anal...This paper explores theoretical solutions to the three-dimensional （3D） shock/shock interaction induced by 3D asymmetrical intersecting compression wedges in supersonic inviscid flows. For Mach interactions, an analytical method known as spa- tial-dimension reduction, which transforms the problem of 3D steady shock/shock interaction into a two-dimensional （2D） pseudo-steady problem on cross sections, is used to obtain the solutions in the vicinity of the Mach stem. The theoretical solu- tions include the pressure, temperature, density, Mach number behind the Mach stem, and total pressure recovery coefficient. Numerical simulations are performed to validate the theoretical results. Here, the NND scheme is employed by solving 3D in- viscid Euler equations, and good agreements are obtained. The asymmetry of the flow characteristics induced by the wedge angle and sweep angle are thoroughly discussed.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
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
Effects of Reynolds(Re)number and Schmidt(Sc)number on the flow structures and variable density mixing are numerically investigated through the canonical shock cylindrical bubble interaction.By determining the viscosi...Effects of Reynolds(Re)number and Schmidt(Sc)number on the flow structures and variable density mixing are numerically investigated through the canonical shock cylindrical bubble interaction.By determining the viscosity and diffusivity within a wide range,the controlling parameters,total vortex circulation,and compression rate,are conservative under a broad range of Re and Sc numbers(Re≈10^(3)-10^(5)and Sc≈0.1-5)in the same shock Mach(Ma)number condition(Ma=2.4).As for the Re number effect,the circulation of secondary baroclinic vorticity(SBV),induced by the main vortex centripetal acceleration,is observed to be higher in high Re number and vice versa.Based on the vorticity transport equation decomposition,a growth-inhibition vorticity dynamics balance mechanism is revealed:the vorticity viscous term grows synchronously with baroclinic production to inhibit SBV production in low Re number.By contrast,the viscous term terminates the baroclinic term with a time lag in high Re number,leading to the SBV production.Since the SBV reflects the local stretching enhancement based on the advection-diffusion equation,mixing is influenced by the Sc number in a different behavior if different Re numbers are considered.The time-averaged variable density mixing rate emerges a scaling law with Sc number asχ^(∗)=β·Sc^(−α),where the coefficientβ∼Re−0.2 and the scaling exponentα∼Re−0.385.The understanding of Re number and Sc number effect on variable density mixing provides an opportunity for mixing enhancement from the perspective of designing the viscosity and diffusivity of the fluid mixture.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
Flow separation due to shock wave/boundary layer interaction is dominated in blade passage with supersonic relative incoming flow,which always accompanies aerodynamic performance penalties.A loss reduction method for ...Flow separation due to shock wave/boundary layer interaction is dominated in blade passage with supersonic relative incoming flow,which always accompanies aerodynamic performance penalties.A loss reduction method for smearing the passage shock foot via Shock Control Bump(SCB)located on transonic compressor rotor blade suction side is implemented to shrink the region of boundary layer separation.The curved windward section of SCB with constant adverse pressure gradient is constructed ahead of passage shock-impingement point at design rotor speed of Rotor 37 to get the improved model.Numerical investigations on both two models have been conducted employing Reynolds-Averaged Navier-Stokes(RANS)method to reveal flow physics of SCB.Comparisons and analyses on simulation results have also been carried out,showing that passage shock foot of baseline is replaced with a family of compression waves and a weaker shock foot for moderate adverse pressure gradient as well as suppression of boundary layer separations and secondary flow of low-momentum fluid within boundary layer.It is found that adiabatic efficiency and total pressure ratio of improved blade exceeds those of baseline at 95%-100%design rotor speed,and then slightly worsens with decrease of rotatory speed till both equal below 60%rated speed.The investigated conclusion implies a potential promise for future practical applications of SCB in both transonic and supersonic compressors.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 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
A novel third-order optimized symmetric weighted essentially non-oscillatory(WENO-OS3)scheme is used to simulate the hypersonic shock wave/boundary layer interactions.Firstly,the scheme is presented with the achieveme...A novel third-order optimized symmetric weighted essentially non-oscillatory(WENO-OS3)scheme is used to simulate the hypersonic shock wave/boundary layer interactions.Firstly,the scheme is presented with the achievement of low dissipation in smooth region and robust shock-capturing capabilities in discontinuities.The Maxwell slip boundary conditions are employed to consider the rarefied effect near the surface.Secondly,several validating tests are given to show the good resolution of the WENO-OS3 scheme and the feasibility of the Maxwell slip boundary conditions.Finally,hypersonic flows around the hollow cylinder truncated flare(HCTF)and the25°/55°sharp double cone are studied.Discussions are made on the characteristics of the hypersonic shock wave/boundary layer interactions with and without the consideration of the slip effect.The results indicate that the scheme has a good capability in predicting heat transfer with a high resolution for describing fluid structures.With the slip boundary conditions,the separation region at the corner is smaller and the prediction is more accurate than that with no-slip boundary conditions.展开更多
This study investigates the three-dimensional(3D)effects introduced by the end walls for an aspect ratio of1 in ramp-induced shock wave boundary layer interactions.The simulations are performed using a symmetry bounda...This study investigates the three-dimensional(3D)effects introduced by the end walls for an aspect ratio of1 in ramp-induced shock wave boundary layer interactions.The simulations are performed using a symmetry boundary condition in the spanwise direction at free-stream Mach numbers in 3D.The simulations are performed using an in-house compressible supersonic solver“Open SBLIFVM”.Two free stream Mach numbers 2.5,and3 are used in the current work,and the simulated results are compared with the aspect ratio 1 simulations by Mangalagiri and Jammy.The inflow is initialized with a similarity solution;its Reynolds number based on the boundary layer thickness is adjusted such that the Reynolds number at the start of the ramp is kept at 3×10^(5)for all simulations.From the results,it is evident that the introduction of sidewalls resulted in a shorter centerline separation length when compared with the two-dimensional(2D)simulations.This contradicts the results at Mach 2 by Mangalgiri and Jammy where the vortex observed at Mach 2 in the central separation region disappeared with increasing free-stream Mach number.Additionally,the topology of interaction shifted from owl-like separation of the second kind to the first kind when the freestream Mach number increased from2 to 2.5.It can be concluded that the interaction topology is crucial to the increase or decrease of the central separation length when compared to 2D simulations.展开更多
The interaction of shock waves is investigated for the following nonstrictly hyperbolic system: [GRAPHICS] The interaction of shock waves is complicated, with new types of shock waves, and new singula rities in the de...The interaction of shock waves is investigated for the following nonstrictly hyperbolic system: [GRAPHICS] The interaction of shock waves is complicated, with new types of shock waves, and new singula rities in the dependence of interaction on the relative positions of the three states separated by shock waves. Several ideas are introduced to helo organize and clarify the new phenomena.展开更多
This paper presents briefly the recent progress on study of swept shock wave/boundary layer interactions with emphasis on application of zonal analysis and correlation analysis to them. Based on the zonal analysis an ...This paper presents briefly the recent progress on study of swept shock wave/boundary layer interactions with emphasis on application of zonal analysis and correlation analysis to them. Based on the zonal analysis an overall framework of complicated interaction flow structure including both surface flowfield and space flowfield is discussed. Based on correlation analysis the conical interactions induced by four families of shock wave generators have been discussed in detail. Some control parameter and physical mechanism of conical interaction have been revealed. Finally some aspects of the problem and the prospects for future work are suggested.展开更多
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.
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co-supported by the National Key Research and Development Plan of China(No.2019YFA0405204)the National Natural Science Foundation of China(Nos.12172365,12072353 and 12132017)。
文摘The interactions of oblique/bow shock waves are the key flow phenomena restricting the design and aerothermodynamic performance of high-speed vehicles.Type Ⅲ and Type Ⅳ Shock/Shock Interactions(SSIs)have been extensively investigated,as such interactions can induce abnormal aerodynamic heating problems in hypersonic flows of vehicles.The transition process between these two distinct types of shock/shock interactions remains unclear.In the present study,a subclass of shock/shock interaction configuration is revealed and defined as Type Ⅲa.Type Ⅲa interaction can induce much more severe aerodynamic heating than a Type Ⅳ interaction which was ever reported to be the most serious in literature.The intense aerodynamic heating observed in this configuration highlights a new design point for the thermal protection system of hypersonic vehicles.A secondary Mach interaction between shock waves in the supersonic flow path of a Type Ⅲ configuration is demonstrated to be the primary mechanism for such a subclass of shock/shock interaction configuration.
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.
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.
文摘A shock interaction problem is solved with finite difference methods for a hypersonic fiow of air with chemical reactions. If a body has two concave cornerst a secondary shock is formed in the shock layer and it meets the main shock later.As the two shocks meet, the flow becomes singular at the interaction point, and a new main shock, a contact discontinuity and an expansion wave appear as a result of interaction between the two shocks. Therefore, the problem is very complicated.Using proper combinations of implicit and explicit finite difference schemes according to the property of the equations and the boundary conditions, we compute the fiow behind the interaction point successfully.
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supported by the National Natural Science Foundation of China (Grant No. 11372333)
文摘This paper explores theoretical solutions to the three-dimensional （3D） shock/shock interaction induced by 3D asymmetrical intersecting compression wedges in supersonic inviscid flows. For Mach interactions, an analytical method known as spa- tial-dimension reduction, which transforms the problem of 3D steady shock/shock interaction into a two-dimensional （2D） pseudo-steady problem on cross sections, is used to obtain the solutions in the vicinity of the Mach stem. The theoretical solu- tions include the pressure, temperature, density, Mach number behind the Mach stem, and total pressure recovery coefficient. Numerical simulations are performed to validate the theoretical results. Here, the NND scheme is employed by solving 3D in- viscid Euler equations, and good agreements are obtained. The asymmetry of the flow characteristics induced by the wedge angle and sweep angle are thoroughly discussed.
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？.
基金This work was supported by the National Natural Science Foundation of China(NSFC)(Grant No.91941301)the Key Research and Development Project of Sichuan Province(Grant No.2019ZYZF0002)。
文摘Effects of Reynolds(Re)number and Schmidt(Sc)number on the flow structures and variable density mixing are numerically investigated through the canonical shock cylindrical bubble interaction.By determining the viscosity and diffusivity within a wide range,the controlling parameters,total vortex circulation,and compression rate,are conservative under a broad range of Re and Sc numbers(Re≈10^(3)-10^(5)and Sc≈0.1-5)in the same shock Mach(Ma)number condition(Ma=2.4).As for the Re number effect,the circulation of secondary baroclinic vorticity(SBV),induced by the main vortex centripetal acceleration,is observed to be higher in high Re number and vice versa.Based on the vorticity transport equation decomposition,a growth-inhibition vorticity dynamics balance mechanism is revealed:the vorticity viscous term grows synchronously with baroclinic production to inhibit SBV production in low Re number.By contrast,the viscous term terminates the baroclinic term with a time lag in high Re number,leading to the SBV production.Since the SBV reflects the local stretching enhancement based on the advection-diffusion equation,mixing is influenced by the Sc number in a different behavior if different Re numbers are considered.The time-averaged variable density mixing rate emerges a scaling law with Sc number asχ^(∗)=β·Sc^(−α),where the coefficientβ∼Re−0.2 and the scaling exponentα∼Re−0.385.The understanding of Re number and Sc number effect on variable density mixing provides an opportunity for mixing enhancement from the perspective of designing the viscosity and diffusivity of the fluid mixture.
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.
基金the funding from the National Key Research and Development Program of China(No.2016YFB0901402)the Key Project of National Natural Science Foundation of China(No.51790513)。
文摘Flow separation due to shock wave/boundary layer interaction is dominated in blade passage with supersonic relative incoming flow,which always accompanies aerodynamic performance penalties.A loss reduction method for smearing the passage shock foot via Shock Control Bump(SCB)located on transonic compressor rotor blade suction side is implemented to shrink the region of boundary layer separation.The curved windward section of SCB with constant adverse pressure gradient is constructed ahead of passage shock-impingement point at design rotor speed of Rotor 37 to get the improved model.Numerical investigations on both two models have been conducted employing Reynolds-Averaged Navier-Stokes(RANS)method to reveal flow physics of SCB.Comparisons and analyses on simulation results have also been carried out,showing that passage shock foot of baseline is replaced with a family of compression waves and a weaker shock foot for moderate adverse pressure gradient as well as suppression of boundary layer separations and secondary flow of low-momentum fluid within boundary layer.It is found that adiabatic efficiency and total pressure ratio of improved blade exceeds those of baseline at 95%-100%design rotor speed,and then slightly worsens with decrease of rotatory speed till both equal below 60%rated speed.The investigated conclusion implies a potential promise for future practical applications of SCB in both transonic and supersonic compressors.
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 Key Basic Research and Development Program (No.2014CB744100)
文摘A novel third-order optimized symmetric weighted essentially non-oscillatory(WENO-OS3)scheme is used to simulate the hypersonic shock wave/boundary layer interactions.Firstly,the scheme is presented with the achievement of low dissipation in smooth region and robust shock-capturing capabilities in discontinuities.The Maxwell slip boundary conditions are employed to consider the rarefied effect near the surface.Secondly,several validating tests are given to show the good resolution of the WENO-OS3 scheme and the feasibility of the Maxwell slip boundary conditions.Finally,hypersonic flows around the hollow cylinder truncated flare(HCTF)and the25°/55°sharp double cone are studied.Discussions are made on the characteristics of the hypersonic shock wave/boundary layer interactions with and without the consideration of the slip effect.The results indicate that the scheme has a good capability in predicting heat transfer with a high resolution for describing fluid structures.With the slip boundary conditions,the separation region at the corner is smaller and the prediction is more accurate than that with no-slip boundary conditions.
基金sponsored by the Department of Science and Technology,Science and Engineering Research Board(SERB),Core Research(Grant No.CRG/2020/03859)。
文摘This study investigates the three-dimensional(3D)effects introduced by the end walls for an aspect ratio of1 in ramp-induced shock wave boundary layer interactions.The simulations are performed using a symmetry boundary condition in the spanwise direction at free-stream Mach numbers in 3D.The simulations are performed using an in-house compressible supersonic solver“Open SBLIFVM”.Two free stream Mach numbers 2.5,and3 are used in the current work,and the simulated results are compared with the aspect ratio 1 simulations by Mangalagiri and Jammy.The inflow is initialized with a similarity solution;its Reynolds number based on the boundary layer thickness is adjusted such that the Reynolds number at the start of the ramp is kept at 3×10^(5)for all simulations.From the results,it is evident that the introduction of sidewalls resulted in a shorter centerline separation length when compared with the two-dimensional(2D)simulations.This contradicts the results at Mach 2 by Mangalgiri and Jammy where the vortex observed at Mach 2 in the central separation region disappeared with increasing free-stream Mach number.Additionally,the topology of interaction shifted from owl-like separation of the second kind to the first kind when the freestream Mach number increased from2 to 2.5.It can be concluded that the interaction topology is crucial to the increase or decrease of the central separation length when compared to 2D simulations.
文摘The interaction of shock waves is investigated for the following nonstrictly hyperbolic system: [GRAPHICS] The interaction of shock waves is complicated, with new types of shock waves, and new singula rities in the dependence of interaction on the relative positions of the three states separated by shock waves. Several ideas are introduced to helo organize and clarify the new phenomena.
文摘This paper presents briefly the recent progress on study of swept shock wave/boundary layer interactions with emphasis on application of zonal analysis and correlation analysis to them. Based on the zonal analysis an overall framework of complicated interaction flow structure including both surface flowfield and space flowfield is discussed. Based on correlation analysis the conical interactions induced by four families of shock wave generators have been discussed in detail. Some control parameter and physical mechanism of conical interaction have been revealed. Finally some aspects of the problem and the prospects for future work are suggested.
