Compression Corner Research Articles

Seismic analysis and connection optimization of the side column on modular prefabricated four‐sided connected composite shear wall

SummaryA modular prefabricated four‐side connected composite (MPFC) shear wall that consists of a composite shear wall modular (CSWM) and steel frame boundary elements is proposed. First, the finite element model (FEM) of the MPFC shear wall, which considers a plastic‐damage constitutive model of both concrete and steel materials, is established based on the finite‐element software ABAQUS. Second, the FEM accuracy of the MPFC shear wall is verified by the experimental results of a modular prefabricated two‐sided connected buckling‐restrained (MTB) steel plate shear wall. Third, the seismic performance of the MPFC shear wall is investigated based on the verified FEM. The connection between the CSWM and side column is optimized. Finally, the initial stiffness calculation formula of the MPFC shear wall that considers the impact of that the thickness ratio between the connection steel plate (CSP) and inner steel plate (ISP) is deduced. The results show that the peak bearing capacity, initial stiffness, total energy, and total strain energy of the MPFC shear wall increased by 42.47%, 44.81%, 113.24%, and 58.97%, respectively, compared with those of the MTB steel plate shear wall. The compressive corner damage of the reinforcement concrete faceplate (RCF) of the MPFC shear wall with a side column and CSWM is connected by a middle steel plate is effectively improved. Compared with those of the MPFC shear wall in which the side column and CSWM are connected by bolts, the shearing force, axial force, and bending moment of the side column of the MPFC shear wall in which the side column and CSWM are connected by the middle steel plate are notably decreased by 25.32%, 26.08%, and 39.51%, respectively. The FEM results are compared with the formula calculation results to establish its accuracy in calculating the initial stiffness of the MPFC shear wall.

A note on the appearance of wave-packets in steady-state triple-deck solutions of supersonic flow past a compression corner

This study is an attempt at solving steady-state triple-deck equations using the method of Bos & Ruban (Phil. Trans. R. Soc. Lond.A, vol. 358, issue 1777, 2000, pp. 3063–3073) for supersonic flow past a compression corner. This was motivated by the fact that in their above paper, they show solutions for scale angles up to 8, the highest obtained so far in the literature. However, we encountered a stationary wave-packet at the corner for scale angles 1.82 and 1.96, depending on the values of stretching factors. Our solutions are then compared with the steady-state solutions produced using the method of Logue, Gajjar & Ruban (Phil. Trans. R. Soc.A, vol. 372, issue 2020, 2014, 20130342), which do not show such wave-packets. These wave-packets do not appear to be the result of flow instability, as flow instabilities should only appear with unsteady equations (Cassel, Ruban & Walker,J. Fluid Mech., vol. 300, 1995, pp. 265–285). It is therefore suggested that the method of Bos & Ruban (Phil. Trans. R. Soc. Lond.A, vol. 358, issue 1777, 2000, pp. 3063–3073) produces these spurious wave-packets as a consequence of their numerical method. This has important implications in the interpretation of triple-deck solutions.

High-enthalpy effects on turbulent coherent structures over a curved compression corner

Direct numerical simulation (DNS) is performed to study high-enthalpy effects on a turbulent boundary layer (TBL) over a curved compression corner. The post-shock flow state behind a wedge flying at Mach 20 and at an altitude of 30 km are chosen for the present simulation. The post-shock temperature is 3400 K, which is high enough to trigger chemical non-equilibrium of the air. A low-enthalpy case is used for comparison. The influences on the instantaneous structures of the streamwise velocity, temperature, and oxygen atoms are examined. The results show that the flow structures are similar on an upstream flat plate in both cases, while on a ramp, streaks of streamwise velocity fluctuations in the high-enthalpy case experience stronger shrink compared with that in the low-enthalpy case. Furthermore, streaks of temperature break into smaller ones when dissociation reactions are introduced. Qualitative and quantitative comparisons are made with the low-enthalpy case; performed using two-point streamwise wall-normal correlation, space–time correlation, and by comparing the propagation velocities of the fluctuations. The results of these analyses validate the observations about the instantaneous fluctuations and show that the differences in the propagation velocity are affected by convection effects and chemical reactions, and that the dissociation reactions accelerate the propagation of temperature fluctuations.

Direct numerical simulation of the effects of Reynolds number in Mach 2.9 flows over an expansion–compression corner

In order to investigate the characteristics of shock wave/boundary layer interaction and its variation with Reynolds number in expansion–compression corner flows, direct numerical simulations of Mach 2.9 expansion–compression corner flows at Reδref=12 000,22 000, and 43 000 were carried out. The λ type shock wave structure appears near the compression corner. Compared with the flat-compression corner wave system, the separation shock wave is weaker, and the reattachment shock wave is stronger. With the increase in the Reynolds number, the separation angle of the separation bubble decreases, the separation pressure rise decreases, and the reattachment pressure rise increases. The separation flow has the unsteady characteristics of low frequency, and the main frequency after nondimensionalized with separation length is St = 0.032. The pressure gradient is the main factor affecting the turbulence change at the expansion corner. With the increase in the Reynolds number, the absolute peak value of pressure gradient at the expansion corner increases, but the spatial range of pressure variation decreases. These two trends together lead to the insignificant change of turbulent energy at the expansion corner. At the compression corner, the main reason for the change of turbulent energy is the strong shear above the separation bubble, which shows a monotone change with the Reynolds number. The results show that the Reynolds number obviously affects the near-wall pressure distribution in the shock disturbance region in the expansion–compression corner flow. The turbulent energy distribution at the expansion corner and compression corner is also affected, but the influence laws are different due to different influence mechanisms.

Performance evaluation of different micro vortex generators in controlling a flare-induced shock–boundary layer interaction

This study evaluates the ability of five micro vortex generator (MVG) geometries in attenuating the flow separation induced by an axisymmetric compression corner. Experimental and computational investigations were carried out at Mach 2 on a cone–cylinder–flare model, with a flow deflection angle of 24° at the cylinder/flare juncture. The MVG shapes considered were baseline ramp, trapezoidal ramp, split ramp, thick vanes and ramped vanes. A circumferential array of these MVGs having a device height (h) of 1.4 mm and an inter-device spacing of 10.5 mm (7.5 h) was introduced 50 mm upstream of the compression corner. Streamwise counter-rotating vortex pairs that originated from these devices created alternate bands of upwash and downwash regions in the incoming boundary layer, which resulted in suitable three-dimensional alterations of the separation region’s size. Furthermore, surface streamline visualizations showed that the vortices induced profound topological transformations in the separated flow structure. In the absence of MVGs, spectral analysis of the pressure signal obtained from the separation shock’s intermittent region revealed a relatively broadband dominant frequency range of 0.55 kHz–0.9 kHz. The MVGs did not cause any significant change in the dominant frequency, but made the bandwidth slightly narrower. Among the different MVG designs that were studied, the ramped vanes (RV) induced the most momentum augmentation in the near wall region and thereby caused the maximum downstream shift in the separation shock’s position.

On wave-packets and discontinuities in triple-deck solutions of supersonic separated flows at a compression corner

We discuss a relationship between wave-packets and discontinuities encountered in the numerical solution of unsteady, supersonic triple-deck equations. Large gradients at the corner are found to compromise the stability of the algorithm of Cassel et al. [3] as the scale angle increases. Two numerical methods have been developed to deal with such gradients. The first method is a time-splitting explicit method with a first-order approximation of the pressure derivative (FOM). This improves the stability of the algorithm at the cost of added numerical diffusion. The range of solutions obtained was extended from 3.7 by Cassel et al. [3] to 6.5. The diffusivity of this method is quantified by comparing with the steady state solutions of Korolev et al. [13] and Logue et al. [19]. The appearance of a spike just prior to the second minimum in shear stress also seen by some previous researchers is elucidated. The second method, which is new, is the Haar Wavelet Method (HWM) applied to unsteady triple-deck equations. This method is found to be stable and accurate, but computationally very expensive. Finally, our results are analysed in terms of the maximum pressure gradient prior to reattachment and the second shear stress minimum as defined and discussed in Smith and Khorrami [27].

A modified wall-adapting local eddy-viscosity model for large-eddy simulation of compressible wall-bounded flow

The wall-adapting local eddy-viscosity (WALE) model in large-eddy simulation can well predict wall-bounded flows but it is also well known for excessive dissipation. In this study, we apply the minimum-dissipation model to constrain the WALE model in compressible flows and obtain the coefficient of the WALE model. Through this process, the dissipation of WALE model can be lower while it still maintains strong stability. In the modified WALE model, the isotropic part of the subgrid-scale (SGS) stress is also reconstructed. In the filtered total energy equation, all of the extra SGS unclosed terms (besides SGS stress and SGS heat flux) are modeled instead of neglecting some SGS terms, such as the SGS viscous diffusion. The modified WALE model is tested in a compressible turbulent channel flow and a supersonic turbulent boundary layer over a compression corner. The new model can well predict the mean velocity, the mean temperature, the Reynolds stress, and the separation bubble.

Bayesian parameter estimation of SST model for shock wave-boundary layer interaction flows with different strengths

The Shock Wave-Boundary Layer Interaction (SWBLI) flow generated by compression corner widely occurs in engineering. As one of the primary methods in engineering, the Reynolds Averaged Navier-Stokes (RANS) methods usually cannot correctly predict strong SWBLI flows. In addition to the defects of the eddy viscosity assumption, the uncertainty of the closure coefficients in RANS models often significantly impacts the simulation results. This study performs parametric sensitivity analysis and Bayesian calibration on the closure coefficients of the Menter k-ω Shear-Stress Transport (SST) model based on the SWBLI with different strengths. Firstly, the parametric sensitivity on prediction results is analyzed using the Sobol index. The results indicate that the Sobol indices of wall pressure and skin friction exhibited opposite fluctuation trends with the increase of SWBLI strength. Then, the Bayesian uncertainty quantification method is adopted to obtain the posterior probability distributions and Maximum A Posteriori (MAP) estimates of the closure coefficients and the posterior uncertainty of the Quantities of Interests (QoIs). The results indicate that the prediction ability for strong SWBLI of the SST model is significantly improved by using the MAP estimates, and the relative errors of QoIs are reduced dramatically.

Separation and flow unsteadiness control in a compression corner induced interaction using mechanical vortex generators: Effects of vane size and inter-device spacing

An experimental investigation was conducted to control separation characteristics of a 24° compression corner induced interaction in a Mach 2.0 flow using an array of mechanical vortex generators (VGs) with rectangular vanes (RRV) placed 6.8δ upstream of the interaction. The objective was to study the effect of (i) inter-VG spacing (s/h = 12, 9.5, 8.0, 6.1, 5.7, 5.5, 4.9, and 4.7), (ii) vane chord length (c/h = 7.2, 4.2, and 3.0), and (iii) vane angle (α = 24°, 20°, 18°, and 16°) in controlling the interaction and on the surface flow topology. These modifications reduce the projected area of VGs in the array from the conventional VG design of RRV2 (c/h = 7.2 and s/h = 9.5) to RRV8 (c/h = 3.0 and s/h = 4.7) by 41%. Reducing s/h also reduces the inter-VG region of the separation significantly that helps to achieve maximum reduction in the streamwise extent of separation up to 83% and in the peak rms value up to 80%. The former improves the overall pressure recovery from 3.0 to 3.4, thereby moving closer toward the inviscid value of 3.8. Surface flow topology shows that the VG array splits a single large spanwise separation bubble for no control into multiple smaller scale individual separation cells placed side-by-side all along the span of the interaction. This helps to reduce the magnitude of mass exchange imbalance across each individual separation cell and, hence, stabilizes the overall interaction relative to no control. The best VG configuration of RRV8 shifts the dominant frequency of fluctuations to approximately 2 kHz or St = 0.19, which is nearly an order of magnitude higher than that for no control.

Direct numerical simulation of shock-wave/boundary layer interaction controlled with convergent–divergent riblets

In this paper, a section of convergent–divergent (C–D) riblets is applied upstream of a compression ramp in a supersonic turbulent boundary layer in a Mach 2.9 flow with a Reynolds number of Reθ=2240. Direct numerical simulations are undertaken to examine the impact of C–D riblets on the shock wave/boundary layer interaction and the feasibility of using them to mitigate flow separation. Over the riblet section, a large-scale secondary roll mode is produced by C–D riblets with the downwelling motion occurring around the diverging region and upwelling motion near the converging region. This consequently leads to a spanwise heterogeneity in mean quantities and turbulent structures over the riblet section and also in the interaction zone. Compared with the baseline case, the area of the separation zone for the riblet case experiences a dramatic local reduction of 92% in the diverging region, owing to the downwelling motion that injects the high-momentum fluid toward the wall and the near-wall spanwise velocity that transports the low-momentum fluid away. The enhanced upwelling motion around the converging region induced by C–D riblets, on the one hand, contributes to the decrease of the near-wall momentum and subsequently the increase of the local separation area. On the other hand, the upwelling motion effectively reduces the incoming Mach number upstream of the compression corner. This appears to reduce the strength of the separation shock, leading to a more gradual compression of the incoming flow that helps ease the enlargement of the separation area nearby. Overall, the area of the mean flow separation is reduced by 56%, indicating an effective flow separation control by the C–D riblets.

Control of supersonic compression corner flow using a plasma actuator

The control performance of a pulsed nanosecond dielectric barrier discharge (NSDBD) plasma actuator with varying pulse voltages and locations on a supersonic compression corner is studied using experiments and numerical simulations. The compression corner with a flat plate length of 60 mm and a ramp angle of 10° under laminar flow separation is experimentally investigated in a Ludwieg wind tunnel under a unit Reynolds number of 7.8 × 106 m−1 and Mach number of 4. The plasma actuators are placed either upstream or downstream of the separation point, extending in the spanwise direction. The Schlieren technique is used to visualize the shock wave interaction and estimate the propagation speed of the induced shock by the plasma actuator. For the numerical simulations, a one-zone inhomogeneous phenomenological plasma model is adopted to predict key discharge parameters and simulate the fast-heating region. The results show that the reduction of separation bubble length is up to 17% and 45%, respectively, in the cases of upstream and downstream of the separation point under a high applied voltage of 50 kV. The evolution of the flow structures is examined to reveal the underlying control mechanism. The results indicate that the high-speed external fluid is entrained into the original separation region after NSDBD activation upstream of the separation point, resulting in flow reattachment upstream of the corner. The entrained fluid with high momentum compels the main separation to move downstream, accompanied by the fragmentation of the original shear layer. Furthermore, the suppression of the separation region is more effective when the plasma actuator is installed close to the separation region and in the first 200 μs during one pulse, providing a good suggestion for the actuation frequency and installed location.

Thermochemical nonequilibrium effects on high-enthalpy double-wedge flows

A hypersonic laminar flow over double wedges with a fixed forward angle of 15° and varied aft angles is studied using computational fluid dynamics and global stability analysis (GSA) at a free-stream Mach number of 12.82 and a total enthalpy of 21.77 MJ/kg. The specific total enthalpy is high enough to trigger evident vibrational excitation and air chemistry. To assess the effects of thermal and chemical nonequilibrium, three different thermochemistry models of air are considered, including frozen, thermal nonequilibrium, and thermochemical nonequilibrium gases. Two-dimensional base-flow simulations indicate that the onset of incipient and secondary separation is insensitive to the inclusion of thermochemistry, although the size of the separation region is substantially reduced. GSA is then performed on the base flows and identifies a three-dimensional stationary global instability beyond a critical aft angle, which is also insensitive to thermochemical nonequilibrium. The criterion of the global stability boundary established for the supersonic flow over compression corners in a calorically perfect gas in terms of a scaled deflection angle [Hao et al., “Occurrence of global instability in hypersonic compression corner flow,” J. Fluid Mech. 919, A4 (2021)] is, thus, extended to high-enthalpy conditions.

Control of flow separation over an axisymmetric flared body using ramped vanes

Experiments were performed at Mach 2 to assess the effectiveness of Ramped Vanes (RVs) in delaying the onset of shock-induced flow separation due to an axisymmetric compression corner. A cone–cylinder-flare model with a 24° flow deflection angle at the cylinder/flare juncture was selected for this study. A single row of RVs was placed 50 mm (10δ) upstream of the compression corner, to manipulate the incoming boundary layer and thereby bring alterations to the shock–boundary layer interaction flowfield. RVs with three sub-boundary layer heights viz. 1.4 mm, 2.1 mm and 2.8 mm (h/δRV=0.33, 0.50 and 0.67 respectively) were employed and their separation control performance was compared. Counter-rotating vortices originated from these devices, generating alternate upwash and downwash bands within the boundary layer. When the shorter devices were used, early and delayed onset of separation were noticed along the upwash and downwash regions respectively. However, as the devices grew taller, the increase in separation length along the upwash region seemed to diminish. Ultimately, when the tallest devices (h=2.8 mm) were introduced, the entire corrugated separation line was successfully pushed downstream of the uncontrolled separation line. Besides instigating these variations in the separation scale, the vortices also caused topological changes within the separation region. But, despite achieving a substantial delay in separation, the RVs were unable to shift the separation shock’s oscillations to higher frequencies.

Unlabeled sample compression schemes and corner peelings for ample and maximum classes

We examine connections between combinatorial notions that arise in machine learning and topological notions in cubical/simplicial geometry. These connections enable to export results from geometry to machine learning. Our first main result is based on a geometric construction by Tracy Hall (2004) [20] of a partial shelling of the cross-polytope which can not be extended. From it, we derive a maximum class of VC dimension 3 without corners. This refutes several previous works in machine learning. In particular, it implies that the previous constructions of optimal unlabeled sample compression schemes for maximum classes are erroneous. On the positive side we present a new construction of an optimal unlabeled sample compression scheme for maximum classes. We leave as open whether our unlabeled sample compression scheme extends to ample classes, which generalize maximum classes. Towards resolving this question, we provide a geometric characterization in terms of unique sink orientations of the associated 1-inclusion graph.

Effect of thermal nonequilibrium on the shock interaction mechanism for carbon dioxide mixtures on double-wedge geometries

The effect of thermal nonequilibrium on shock interactions of carbon dioxide (CO2) hypersonic flows is investigated. Given the relatively low characteristic vibrational temperature of the CO2 molecule, it is expected that excited vibrational modes play a significant role in the physics of shock/shock and shock/boundary layer interactions. The shock interference mechanism resulting from a CO2-dominated flow over different double-wedge geometries is investigated by numerically solving the Navier–Stokes equations within the framework of a two-temperature model that considers translational energy–vibrational energy transfer. To assess the impact of vibrational relaxation, a comparative assessment of the patterns obtained with three thermo-physical models is presented, with the two-temperature model flow pattern being compared to thermally perfect and perfect ideal gas ones. Results obtained with the two-temperature model show that increasing the aft angle significantly enlarges the separated region in the compression corner and generates numerous secondary shock waves and shear layers. Peaks of heat flux and pressure occur along the surface due to boundary layer reattachment downstream of the compression corner, except for the case of the higher angle, which results in the largest peaks due to shock impingement. Different assumptions on the excitation of vibrational modes are shown to largely influence the size of the recirculation bubble in the compression corner, shock interaction mechanism, and surface loads. The more energy transferred to the vibrational mode, the lower post-shock temperatures are obtained, which tends to reduce the post-shock density, leading to weaker shock interactions characterized by delayed onsets of separation, reduced separation regions, and smaller standoff distances.

Three-dimensionality of hypersonic laminar flow over a double cone

Hypersonic laminar flow over a canonical 25°–55° double cone is studied using computational fluid dynamics and global stability analysis (GSA) with a free-stream Mach number of 11.5 and various unit Reynolds numbers. Axisymmetric simulations reveal that secondary separation occurs beneath the primary separation bubble beyond a critical Reynolds number. The numerical results agree well with existing experiments and the triple-deck theory with the axisymmetric effect on the incoming boundary layer treated by the Mangler transformation. The GSA identifies a three-dimensional global instability that is azimuthally periodic immediately prior to the emergence of secondary separation. The criterion of the onset of global instability in terms of a scaled deflection angle established for supersonic compression corner flows (Hao et al., J. Fluid Mech., vol. 919, 2021, A4) can be directly applied to double-cone flows. As the Reynolds number is further increased, the flow is strongly destabilized with the coexistence of multiple stationary and low-frequency oscillating unstable modes. Direct numerical simulations confirm that the supercritical double-cone flow is intrinsically three-dimensional, unsteady and exhibits strong azimuthal variations in the peak heating.

Numerical investigation of a Mach 6 hypersonic laminar flow on two-dimensional cold-wall compression corners with controlled surface roughness

Numerical simulations are performed in order to investigate the physics of a Mach 6 hypersonic, laminar flow on different configurations of a two-dimensional compression corner. For this geometry, typically characterising control surfaces of re-entry vehicles, the interaction between the incoming boundary layer and the ramp-induced shock wave can lead to the formation of a separated-flow region. The numerical experiments, carried out under adiabatic and a cold-wall conditions, include the presence of a two-dimensional sinusoidal roughness patch, which is placed at various locations upstream of the ramp, in order to evaluate the effects of the surface waviness on the separation bubble. When the roughness elements are placed upstream of the smooth separation point, the recirculation bubble stretches towards the roughness position. This behaviour is induced by the thickening of the boundary layer downstream of the roughness patches, which cause a displacement of the laminar flow. No transition is observed downstream of the patches under the investigated geometrical and flow conditions. For a patch downstream of the smooth separation point, instead, no appreciable effects on the bubble length are noticed. However, the surface roughness inside the bubble can alter the recirculation pattern within the reverse-flow region. Hence, together with the length of the separated flow region, also the heat transfer at the wall is affected by the roughness because of the different recirculation patterns inside the bubble. In particular, the Stanton number highlights a redistribution of the heat-flux intensity between the main peak at reattachment and the secondary peaks inside the separated flow. This effect is more evident for higher ramp angles. These results highlight the importance of roughness-induced effects on the complex phenomenon that shock-wave/boundary-layer interaction is.

High-Order Spectral/-Based Solver for Compressible Navier–Stokes Equations

High-order spectral element methods are becoming more established for solving problems in computational fluid dynamics. In this paper, we introduce new stabilization parameters for the solution of compressible Navier–Stokes equations in conjunction with high-order spectral element methods. We previously defined a new set of stabilization parameters for Euler equations in a high-order spectral/ framework. In this paper, we examine the definition of these stabilization parameters in the context of solutions of full Navier–Stokes equations, and we report good agreement with previously published results in the literature. We establish spectral convergence of errors for Kovasznay flow. We then validate the definition of the stabilization parameter for Couette flow, flat plate, and compression corner problems. In a later example, we solve the flow past a cylinder and benchmark results with previously published results obtained with low-order methods and other approaches in compressible flow computations. We further solve hypersonic flow past a wedge and supersonic flow past a Flight Investigation of Re-Entry program (FIRE) entry capsule. We explore the definition of the shock-capturing parameter based on the parameter. The introduced definitions are found to provide excellent results for the full Navier–Stokes equations.

Mechanism analysis of magnetohydrodynamic control in hypersonic turbulent flow

Under the assumption of the low magnetic Reynolds number, the coupled model is established for the turbulent flow field and the externally applied magnetic field. The AUSMPW+ scheme and LUSGS method are used to solve turbulent magnetohydrodynamics (MHD) flow equations, in which the Spalart-Allmaras one-equation turbulence model is used. A series of numerical simulations over various geometry configurations, namely, a flat plate and a compression corner, is conducted by using an external electromagnetic field. Results show that the performance of MHD boundary layer flow control is determined mainly by the Lorentz force in the streamwise direction. With an external magnetic field used, the low velocity fluid in the boundary layer can decelerate and increase the static temperature locally. Moreover, the counter-flow Lorentz force always brings a negative effect on the turbulent skin friction coefficient, and the location for the MHD zone has a great influence on the control efficiency of the ramp-induced separation. A reasonable magnetic field layout scheme should be configured in practical engineering application.

Thermochemical non-equilibrium effects on hypersonic shock wave/turbulent boundary-layer interaction

The research on shock wave/turbulent boundary-layer interactions is mainly limited to calorically perfect gases; little has been reported on the thermochemical non-equilibrium (real gas) effect. This effect is prominent at conditions of high Mach and Reynolds numbers. In this work, a household parallel solver for hypersonic thermochemical non-equilibrium flows with a Reynolds-averaged Navier–Stokes turbulence model is developed in which the coupling of turbulence with vibration and chemistry occurs under a gradient-law assumption. The thermal non-equilibrium is based on Park's two-temperature model, and the chemical non-equilibrium is based on Gupta's 11-species model. The method proposed in this paper is first validated using experimental data, including cases of a laminar cylinder flow at a high-enthalpy condition, a supersonic flat-plate turbulent boundary layer flow, a hypersonic transition flow, and a hypersonic compression corner flow at a low-enthalpy condition. This approach is then applied to assess the hypersonic flow characteristics past the 34° compression corner at a flight height of 30 km. Results show that the joint effects of turbulence and thermochemical non-equilibrium have a significant impact on the flow field organization, wall data, and separation length of the shock wave/boundary-layer interaction. Furthermore, the mechanism of the neck region accompanied by maximum heat flux, wall pressure and skin friction in both laminar and turbulent cases is well-interpreted. This study can be used as a reference tool for the aerodynamic design of future hypersonic vehicles accounting for multi-physics effects.