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In this paper, by using a large eddy simulation, we study the combustion process in the HyShot II scramjet combustor. By conducting a detailed analysis of the mass-fraction distributions of the main species, such as H 2 , H 2 O and the radicals, OH and HO 2 , of the mass source terms of these main species, and of the chemical source term of the energy equation, we detect the regions where chemical reactions occur through a diffusion process and the regions where auto-ignition and premixed combustion may develop. The analysis indicates that the combustion process is mainly of diffusive type, along a thin shear layer enveloping the hydrogen plume, whereas, there could be some auto-ignition and/or premixed combustion cores inside the plume. • An innovative propulsion system for hypersonic flight was analysed. • A large eddy simulation approach was employed to study the HyShot II combustor. • A thorough analysis of both fluid dynamic and chemical aspects was performed. • More understanding of the aerothermochemistry phenomena is provided. [ABSTRACT FROM AUTHOR]

This paper takes a new look at how the aerodynamic interactions of multiple bodies in high-speed flow affect their motion behaviors. The influence of the body shape and orientation on aerodynamic and stability behavior in the case of shock–shock and wake–shock interactions is the focus of this publication. Experiments were performed in the hypersonic wind tunnel H2K at the German Aerospace Center (DLR) in Cologne. Free-flight tests with tandem arrangements of spheres and cubes were performed with a synchronized dropping of both objects at various initial conditions of relative streamwise and vertical distance as well as pitch angle. A high-speed stereo-tracking captured the model motions during free-flight, and high-speed schlieren videography provided documentation of the flow topology. Based on the measured 6-degrees-of-freedom (6DoF) motion data, aerodynamic coefficients were determined. As a result, the final lateral velocity of trailing cubes is found to be many times greater than that of spheres regarding shock-wave surfing. For rotating cubes, the results showed that stable shock-wave surfing can become possible over an increasingly wide range of initial positions. This study has identified that the trailing drag coefficient of two axially aligned objects varies strongly with their relative streamwise distance. Furthermore, it was shown that the wake is a region of stability for downstream objects. [ABSTRACT FROM AUTHOR]

In this study, we propose a scheme to control the deflection of the warhead based on the configuration of the hypersonic glide body (HGB) to solve the problems posed by its large control surface load and severe aerodynamic heat under hypersonic flight conditions. We conducted numerical simulations on the configurations of deflection of the warhead of an HGB analog under different flight modes as well as varying angles and directions of deflection. The results showed that once the warhead had been deflected, the overall configuration of the HGB analog still exhibited static longitudinal stability. An increase in the angle of deflection significantly reduced the lift-to-drag ratio of the configuration at large angles of attack. When the warhead was deflected upward, the configuration of the HGB analog exhibited static lateral instability, while it exhibited a high static lateral stability when the warhead was deflected downward. [ABSTRACT FROM AUTHOR]

The current work is undertaken to examine the capability of a recently developed ICFM (internal conical flow M) basic flowfield [Musa et al., AIAA J. 61, 1–16 (2023)] to design hypersonic internal waverider intakes. The osculating axisymmetric flows concept is employed alongside the streamline-tracing method to design three internal waverider intakes with different contraction ratios, i.e., 6.0, 9.0, and 25. The three intakes have similar capture and throat shapes and entrance areas. Then, the starting characteristics of designed intakes were investigated at different design and off-design conditions. The selected design condition is the flight at Mach 6.0 and 28 km altitude. Viscous computations have been performed at Mach 6.0, 5.0, 4.0, 3.9, 3.8, 3.5, and 3.0 for the intake with a contraction ratio of 6.0, and at Mach 6.0, 5.0, 4.7, and 4.0 for the intake with contraction ratio of 9.0. The intake with a contraction ratio of 25 is simulated at Mach 6.0 and 5.0. Successful intake starting has been achieved up to Mach 3.8 for the first intake and up to Mach 4.7 for the second intake. The intake with a contraction ratio of 25 attained an unstart state in the design condition due to a very high contraction ratio. Thus, the starting limit of internal waverider intakes designed using ICFM basic flowfield is found to be controlled by Van Wie empirical limit. It is concluded that the ICFM basic flowfield can effectively design high-performance fixed-geometry intakes with overboard spillage, demonstrating excellent performance and startability across a wide range of Mach numbers. [ABSTRACT FROM AUTHOR]

In this work, the rarefied Couette flow of diatomic gases with thermal nonequilibrium effects is investigated by the direct simulation Monte Carlo (DSMC) method, and a macroscopic computational model is developed to consider the local rarefaction effects for diatomic gases in the near-continuum regime. The nonlinear transport properties of the diatomic gases are studied, indicating that effective viscosity and effective translational thermal conductivity in the shear nonequilibrium state are affected by translational nonequilibrium effects, which obey the same laws for both monatomic and diatomic gases. The transport coefficients of internal energy modes are affected by both translational nonequilibrium and internal energy relaxation, therefore, the effective rotational and vibrational thermal conductivities are related to the effective viscosity through a modified Eucken relation that accounts for internal energy relaxation. Conclusively, effective constitutive relations are newly established as a function of the shear nonequilibrium parameter and the modified Eucken factors for thermal nonequilibrium flows, and these are integrated into the macroscopic two-temperature model. Subsequently, it is assessed in the simulation of hypersonic flows over flat plates and cylinders at various Knudsen numbers. The results show that the surface shear stress and heat flux obtained by the proposed model agree well with the DSMC results, indicating significantly improved performance compared to the conventional Navier–Stokes two-temperature model for hypersonic flows in the near-continuum regime. [ABSTRACT FROM AUTHOR]

In nonequilibrium slip and jump conditions, normal gas velocity and temperature gradients are used to calculate the gas slip velocity and temperature at the surface, respectively. Gökçen et al. (Computational fluid dynamics near the continuum limit, AIAA Paper No. 87-1115, 1987, and Gökçen and MacCormack, Nonequilibrium effects for hypersonic transitional flows using continuum approach, AIAA Paper No. 89-0461, 1989) stated that the tangential velocity and temperature of the gas molecules before a collision with the surface could be interpreted as the macroscopic tangential velocity and temperature of the gas molecules at the so-called near-wall distances auλ and aTλT away from the surface, respectively. The coefficients au and aT are the order of unity. In the present work, new forms of the slip and jump conditions are proposed by modifying the Gökçen slip and jump conditions to include the coefficients (au, aT). Numerical investigations are comprehensively conducted to determine the numerically proper values (au, aT) for the hypersonic rarefied gas flows. Cases such as the circular cylinder in cross-flow and sharp and blunted leading edge wedge are considered in the present work, with nitrogen as the working gas. The simulation results show the significant effects of the coefficients (au, aT) on the accuracy of the slip velocity and surface gas temperature predictions, and the values of au = 1.2 and aT = 1.1 show good agreement with the direct simulation Monte Carlo data. [ABSTRACT FROM AUTHOR]

High Reynolds number turbulent flow of hypersonic vehicles exhibits multi-scale flow structures and non-equilibrium high-frequency characteristics, presenting a significant challenge for accurate prediction. A deep neural network integrated with attention mechanism as a reduced order model for hypersonic turbulent flow is proposed, which is capable of capturing spatiotemporal characteristics from high-dimensional numerical turbulent data directly. The network model leverages encoder–decoder architecture where the encoder captures high-level semantic information of input flow field, Convolutional Long Short-Term Memory network learns low-dimensional characteristic evolution, and the decoder generates pixel-level multi-channel flow field information. Additionally, skip connection structure is introduced at the decoding stage to enhance feature fusion while incorporating Dual-Attention-Block that automatically adjusts weights to capture spatial imbalances in turbulence distribution. Through evaluating the time generalization ability, the neural network effectively learns the evolution of multi-scale high-frequency turbulence characteristics. It enables rapid prediction of high Reynolds number turbulence evolution over time with reasonable accuracy while maintaining excellent computational efficiency. [ABSTRACT FROM AUTHOR]

During the flight of hypersonic vehicle, air will be decomposed and ionized due to the "friction" under ultra-high speed, thus forming a plasma layer. Because the plasma has the ability to absorb and reflect electromagnetic waves, the "black barrier" phenomenon is formed. In addition, when the hypersonic vehicle passes through the atmosphere, the surface temperature rises sharply due to aerodynamic heating, and the surface material undergoes a series of complex changes to form ablation. In this paper, the finite volume method and the laminar finite rate model are used to study the flow field velocity, pressure distribution, flow field temperature, and spatial distribution of each component of the aircraft at different Mach numbers, angles of attack, and heights. In the flow field of supersonic aircraft, N and O are mainly concentrated in the tail of the aircraft, NO is mainly concentrated in the head of the aircraft, and N2 and O2 are full of the whole space. Because of the accumulation of NO+ and O2+ in the tail of the aircraft, the charge accumulation is formed, which will further interfere with the electromagnetic wave signal. The mass fraction of N and O increases with the increase in Mach numbers, while the mass fraction of O2 decreases with the increase in Mach numbers. Different angles of attack will affect the asymmetry of the shock wave of the aircraft. In this paper, the evolution mechanism and distribution characteristics of aircraft plasma are revealed, which lay a theoretical foundation for solving the problem of black barrier and ablation. [ABSTRACT FROM AUTHOR]

Gaps represent a common feature on the surface of supersonic vehicles, and their presence engenders localized thermal phenomena, including elevated heat flux and thermal invasion. These occurrences significantly imperil the aircraft's thermal stability and safety. This study employs numerical simulation to investigate supersonic streamwise gap flow and offers a comprehensive analysis of the impact of geometric and leakage on the distribution of wall heat flux. The computational results manifest a high alignment between regions of high heat flux and high Mach number. Consequently, these findings underscore the pivotal role of convective heat transfer governing the distribution of heat flux within the gap. Most notably, the sensitivity of wall heat flux is most pronounced with respect to gap width. A reduction in gap width yields a substantial reduction in wall heat flux, with this alteration primarily stemming from its impact on the shear interactions between the supersonic mainstream and the vortices within the gaps. Furthermore, changes in gap length influence the invasion characteristics of the supersonic mainstream, thereby instigating consequential shifts in the distribution of heat flux density along the inner surfaces of the gaps. In contrast, variations in gap depth exert a comparatively modest impact on the distribution of heat flux density. The introduction of leakage effects at the base of the gaps results in a significant intensification of supersonic mainstream penetration and concomitant convective heat transfer in the bottom exit region. This phenomenon engenders localized zones characterized by elevated heat flux density, thus severely compromising the aircraft's thermal sealing performance. • Thermal environments within streamwise gaps in supersonic flow are investigated. • Influences of geometries on the distribution of wall heat flux are analyzed. • Mechanisms by which changes in vortex affect the thermal environment are revealed. • Effects of potential bottom leakage on thermal invasion are evaluated. [ABSTRACT FROM AUTHOR]

The curved surfaces widely exist in three-dimensional inward-turning inlets and variable sectional isolators, which determine the downstream evolution of the converged flow generated by the crossing shock wave/boundary layer interaction, thereby affecting the start performance and combustion efficiency of the hypersonic vehicles. In this work, the converged flow subjected to the curved surfaces is experimentally studied with the nanoparticle-based planar laser scattering method and particle image velocimetry. Results show that the velocity profiles of the converged flow on the transverse plane can be normalized at different angles. The upstream separation size is found to be decreased by the downstream convex surface due to the reduction of backpressure, especially in large-radius cases, whereas the converged flow gradually becomes smooth ought to the dependence of the acceleration effect on the velocity. The concave surface shows an opposite effect on the converged flow, which further indicates that a downstream convex surface is preferred for hypersonic inlets to weakens the flow non-uniformity to improve the inlet performance. • Converged flow generated by the crossing shock wave/boundary layer interaction on the transverse plane is visualized semi-elliptical. • Velocity profile of the converged flow on the transverse plane can be normalized at different angles and expressed in the logarithmic form. • The convex surface makes the converged flow dissipated under the momentum exchange and the expansion effect. • A downstream convex surface is preferred for inlet design. [ABSTRACT FROM AUTHOR]

Scramjet engines represent one of the most promising alternatives to power hypersonic vehicles, although the peculiarities of supersonic combustion raise many challenges. Computational Fluid Dynamics (CFD) simulations can be used to model highly turbulent reactive flows in such complex configurations and avoid expensive experimental campaigns. A particularly challenging aspect when dealing with numerical simulations of turbulent combustion is accurately modelling the interactions between chemistry and turbulence. These interactions can play a significant role in supersonic flows and require an appropriate closure since the conditions encountered in supersonic flows differ from those of conventional combustion devices. For this reason, an assessment of the role of turbulence-chemistry interactions in supersonic combustion is needed. This work evaluates the performance of the Partially-Stirred Reactor (PaSR) closure of the mean chemical source term by performing Reynolds-Averaged Navier–Stokes (RANS) simulations of two selected test cases, namely, the Burrows and Kurkov combustor and the German Aerospace Centre (DLR) scramjet. The PaSR approach requires the definitions of two important quantities: mixing and chemical time scales. While the chemical time scale is related to finite-rate chemistry, supersonic flow conditions can affect the most appropriate definition of the mixing time scale. For this reason, several formulations of the mixing time scale are tested in this work. The PaSR model is validated against experimental data and also compared to the finite-rate approach without turbulence-chemistry interactions (FR/No-TCI) and the Eddy Dissipation Concept (EDC). In the Burrows and Kurkov test case, the PaSR model can accurately predict the main quantities of interest (e.g., temperature and chemical compositions), showing better results than the FR/No-TCI and similar to the EDC model. The validation on the DLR case confirms these trends and highlights the importance of accounting for turbulence-chemistry interaction in supersonic combustion modelling. However, the EDC model is not able to predict a stable ignited flame in the DLR combustor. The presence of recirculation zones near the strut injector of the DLR scramjet can explain the better performance of the PaSR model versions employing the integral-based and geometric mean formulations, which estimate comparable values of the mixing time but higher than those of other time scale approaches considered in this work. • Supersonic combustion is characterised by similar chemical and mixing time scales. • The Partially Stirred Reactor model is suitable for RANS simulations of supersonic combustion. • The model performs better than a laminar finite-rate approach and Eddy Dissipation Concept. • The choice of the mixing time scale affects the Partially Stirred Reactor accuracy. • The most appropriate mixing time scale is selected for two scramjet combustion test cases. [ABSTRACT FROM AUTHOR]

The current work evaluates the effectiveness of the pressure-corrected osculating axisymmetric flows method in the design of hypersonic wavecatcher intakes with shape transition. The original osculating axisymmetric flows method, which is essentially used in waverider design, has drawbacks in accurately demonstrating the three-dimensional flowfield due to ignoring the cross-flow effects between the osculating planes. This negatively impacts the intake performance due to the presence of lateral pressure gradients. The pressure-corrected method takes into account these effects by incorporating lateral pressure gradient corrections into the original design methodology using cross-flow velocity information. This results in the generation of three-dimensional streamlines rather than the two-dimensional streamlines of the original method. The semirectangular-to-ellipse wavecatcher intake is selected as the subject of investigation, and the design procedure is reviewed. The characteristics of the wavecatcher intakes with a design point of Mach 6.0 are studied. Computational fluid dynamics analysis of pressure-corrected wavecatcher intakes is presented to assess the design technique. It is found that the initial conical shock impinges at the intake's entrance and the streamtube is completely captured. Furthermore, the comparison with the original wavecatcher intake indicates that the pressure-corrected wavecatcher intake demonstrated better performance in terms of total pressure recovery and flow uniformity. The wavecatcher intake with entrance-to-exit shapes transition showed higher performance than those with the same entrance and exit shapes. [ABSTRACT FROM AUTHOR]

High-speed flows over a 34° compression corner subjected to several types of applied electromagnetic fields are numerically investigated. A coupled model for the turbulent flow field and the externally applied electromagnetic field is established based on the low magnetic Reynolds number assumption. The advection upstream splitting method with pressure-based weight function (AUSMPW) scheme and the lower upper symmetric Gauss Seidel (LUSGS) method are implemented to solve turbulent magnetohydrodynamic (MHD) flow equations. Numerical results demonstrate that the performance of MHD separation flow control is determined mainly by the Lorentz force in the turbulent boundary layer. For the counter-flow Lorentz force, the velocity fluid in the boundary layer can be retarded, and a significant temperature increase was observed. The decelerating Lorentz force significantly brings negative effects on the turbulent skin friction coefficients and increases the static pressure locally. Furthermore, the position of the MHD zone has a significant impact on the control efficiency of the ramp-induced separation. With an external electric and magnetic field applied, the low-velocity fluid in the boundary layer can be accelerated, and the Lorentz force directed along the stream can reduce the size of the separation bubble. [ABSTRACT FROM AUTHOR]

Hypersonic air-breathing propulsion (>Mach 5) based on scramjet offers new solutions to both the military strike platform and the civil transportation. However, the hypersonic flight seriously challenges the engine thermal protection. One of the key problems is the tricky flow and thermal behaviors of hydrocarbon fuel in regenerative cooling channels under the complex thermal boundaries and channel geometry, which may cause heat transfer deterioration and structure over-temperature. To improve the cooling design, the double-layer regenerative cooling channels are focused and compared with the traditional single-layer channels. The flow and thermal behaviors of hydrocarbon fuel under different cooling configurations are numerically studied. Cooling performance of double-layer channels is superior than single-layer channels. Tw,max falls by 295.2 K/17.7%. Gravity direction and magnitude show negligible influence on the heat transfer. Regarding channel height ratio, H1/H2 = 1.1/1 achieves optimal cooling performance in present conditions. Double-layer-channel scheme presents high practicability in scramjet cooling. [ABSTRACT FROM AUTHOR]

High flight speed of hypersonic vehicles will push thermal insulation composites to achieve balanced enhancement of high porosity with increasing load-bearing capability. Herein, hierarchically porous SiC(rGO) polymer-derived ceramics (PDCs) with lightweight, heat-insulating, anti-ablative and tough engineered merits, are designed from polycarbosilane-vinyltriethoxysilane-graphene oxide precursors by re-pyrolysis/graphite-assisted decarburization. Graphite pore-forming agents tightly coupled with SiO x C y /C free (rGO) by C-dangling bonds, can inhibit SiO x C y decomposition, increase ceramic network disorder and expedite concurrent removal of C free. SiO 2 generated from Si-dangling oxidation bonding, efficiently densifies framework to retain mechanical performances, and meanwhile creates more interfaces to enhance phonon scattering. Such structure similar to nodular cast iron confers low thermal conductivity (0.11 W·m-1·K-1), particularly, fully preserves good compressive strength (5.4 MPa) for SiC(rGO) 40%. Under both cyclic butane flame ablation (∼1300 °C) and long-term oxidation for 7200 s, self-healing effect of SiO 2 imbues products with good structural integrity, shedding light on the opportunity to apply any thermal protection system. [Display omitted] • Hierarchically porous SiC(rGO) PDCs were ingeniously in-situ constructed. • Re-pyrolysis/graphite-assisted decarburization is a new structural design strategy. • The obtained products own well-balance between thermal insulation and load bearing. • Graphite-assisted interfacial interactions aims at strengthening ceramic framework. • Self-healing effect of SiO 2 can be promptly activated under flame scouring. [ABSTRACT FROM AUTHOR]

In this paper, a one-dimensional (1-D) range profile of the hypersonic target enveloped by a plasma sheath is investigated. Firstly, the non-uniform property of the plasma sheath is studied and its impact on the wideband electromagnetic (EM) wave is analyzed. A wideband radar echo model for the plasma-sheath-enveloped hypersonic target is constructed. Then, by exploiting the relationship among the incident depth, reflection intensity, and plasma velocity, it reveals that distinct scatter points in various areas of the target will suffer from varying reflection intensity and coupled velocity, leading to severe defocusing in the range profile. To tackle this issue, a novel focusing algorithm combing the Fractional Fourier Transform (FRFT) with the CLEAN technique is developed, which independently calculates the coupled plasma velocity and compensates for the phase error via a series of iterative procedures. Finally, the influence of the plasma sheath on the 1-D range profile and the effectiveness of the proposed focusing algorithm are validated through simulations. [ABSTRACT FROM AUTHOR]

Presently, there is significant emphasis on researching the automation of processes, systems, and devices across scientific and technical fields. This emphasis extends notably to sectors like aerospace, robotics, and electric transportation technology. One effective approach to tackling the technical and design challenges of automating various technological tools and processes is through the implementation of adaptive systems. Specifically, achieving efficient and reliable power transmission in electric autonomous and mobile vehicles can be realized by employing an adaptive mechanical drive with two degrees of freedom. The adaptability of this two-mobile system is enabled by introducing an innovative continuous multi-speed drive design featuring an additional friction coupling. The primary attribute ensuring the reliability of this self-adjusting mechanical drive lies in its ability to autonomously adapt to external loads. This is made possible by a balancing friction coupling that establishes a connection between the frictional torque and the relative angular velocity. [ABSTRACT FROM AUTHOR]

Near-space hypersonic vehicles could encounter significant rarefied nonequilibrium effects during the flight through atmosphere, which largely influence the gas-surface momentum and heat transfer. In this paper, hypersonic boundary layer over a flat plate with velocity slip, temperature jump, and shear nonequilibrium effects is theoretically considered. The slip boundary conditions and nonlinear transport relations are embedded into the boundary-layer equations to describe the flow. Local similar solutions are derived, and key parameters for characterizing slip and shear nonequilibrium effects are determined. The velocity-slip and temperature-jump effects are determined by [ (2 − σ u) / σ u ] M a e / R e x and [ (2 − σ T) / σ T ] M a e / R e x respectively, and the shear nonequilibrium effect is characterized by M a e 2 / R e x . The obtained boundary-layer solutions are compared with the Navier–Stokes solutions for a Mach 4.5 slip flow, and the results of Direct Simulation Monte Carlo for a Mach 10 rarefied flow, good agreements are achieved. The separate and combined effects of velocity slip, temperature jump, and shear nonequilibrium on boundary-layer solutions and momentum/heat transfer are clarified. The results show that both the slip and shear nonequilibrium effects cause the boundary layer to become thinner and decrease the skin friction and Fourier heat conduction. However, with including sliding friction, the total heat flux might even increase as the slip degree increases. These results provide valuable insight into the boundary-layer characteristics of hypersonic near-continuum flows. [ABSTRACT FROM AUTHOR]

Transition of hypersonic boundary layer flow is an important problem in aerodynamic research. To obtain the natural transition position using the flow stability analysis method, it is necessary to analysis the neutral curve and the perturbation growth rule. Accurate calculation of the basic flow is the basis and prerequisite for stability analysis. This paper proposes that surface radiation should be considered in the analysis of hypersonic boundary layer flow stability due to the high temperatures at the vehicle surface caused by aerodynamic heating. In order to obtain the influence of surface radiation on the flow stability, we assume that the surface satisfies the energy balance and the gas satisfies calorimetrically perfect gas properties. To simplify the analysis, gas radiation has been neglected. The boundary layer analysis of a hypersonic flat plate with and without radiation has been performed by a combination of flow stability theory and direct numerical simulation. It is found that high temperature radiation on the surface causes energy to transfer outwards. The results showed that surface temperature decreases along the streamwise direction and is no longer an adiabatic surface temperature. Instability interval decreases and the maximum growth rate increases. The range of the instability region expressed in dimensionless frequency values moves to lower frequency as the Mach number increases. Maximum growth rate shows a tendency to increase and then decrease as the Mach number increases. [ABSTRACT FROM AUTHOR]

This paper introduces a novel hybrid numerical method, SAUSM, designed for accurate and robust simulation of compressible flows governed by the Euler equations. While the AUSM + scheme provides proper resolution of smooth flow features, it is susceptible to anomalies, particularly the carbuncle phenomenon near strong shock discontinuities. Conversely, the AUFS scheme offers inherent stability in capturing shocks; however, it lacks the accuracy of AUSM + in smooth regions. The proposed SAUSM method combines AUSM + and AUFS through an adaptive weighting function, facilitating a seamless transition between the schemes. This approach preserves the accuracy of AUSM + in smooth regions while ensuring robust shock-capturing capabilities near discontinuities. The effectiveness of the SAUSM method is rigorously demonstrated through a comprehensive suite of progressively complex test cases. Numerical experiments demonstrate SAUSM's proficiency in resolving intense shock patterns and discontinuities without introducing anomalies. In the selected test cases, SAUSM agrees with reference solutions and effectively mitigates anomalies observed in AUSM + , including kinked Mach stems. In the challenging test case involving hypersonic blunt body flow over a cylinder, SAUSM adapts dissipation effectively by utilizing its adaptive weighting function to generate smooth pressure distributions, thereby eliminating the carbuncle instability linked to AUSM + when applied to a high aspect ratio grid. The consistent formulation of flux splitting and the adaptive weighting in SAUSM prevent excessive dissipation away from discontinuities, thus preserving accuracy comparable to that of exact Riemann solvers. Consequently, SAUSM emerges as a promising and innovative approach to accurately and robustly simulate a wide range of compressible Euler flows. The comprehensive results obtained from the validation tests firmly establish SAUSM as a highly effective general-purpose technique for computational fluid dynamics in academic research. [ABSTRACT FROM AUTHOR]

Quantitative characterization of freestream fluctuations in a hypersonic flow facility is important for proper interpretation of results from any boundary-layer flow transition experiments conducted in the facility. Direct measurement of freestream flow quantities (like pressure, temperature, velocity) using a physical probe is not possible due to the modulation introduced by the leading shock wave and the downstream flow region. A transfer function that models the complex modulation process, which is specific to a given probe geometry, is therefore needed to estimate the freestream flow condition from the physical probe output. The general practice across different experimental hypersonic flow facilities globally is to employ custom designed probes for estimation of freestream fluctuations. This hinders a direct comparison of freestream fluctuations between two different flow facilities since such a comparison is limited by the accuracy of the two different transfer functions that are used. The present work is motivated by the need to have a standard probe geometry and an associated transfer function for enabling an objective comparison of freestream fluctuations across experimental facilities. A wedge probe along with a linear transfer function used in the recent work by Wagner et al. (J Fluid Mech 842:495–531, 2018. https://doi.org/10.1017/jfm.2018.132) is proposed for consideration as a standard hypersonic freestream flow characterization method. The utility of the measurement method is demonstrated by making spectral measurements of freestream pressure fluctuations at Mach 6 in the Roddam Narasimha Hypersonic Wind Tunnel at IISc and performing a direct quantitative comparison with the results obtained by Wagner et al. in the RWG Ludwieg tube (Göttingen, Germany). [ABSTRACT FROM AUTHOR]

Conceptual design studies performed with reduced-order approaches allow feasibility and sizing considerations of air-breathing engines to be configured at an affordable computational cost. The present study is devoted to exploring the design space of a dual-mode ramjet engine operating in scramjet mode by means of reduced-order analysis to assess the effects of propulsive system design configurations on component level and overall performance characteristics. The approach proposed in this work combines axisymmetric flow configuration used for the design of supersonic/hypersonic intakes and solutions of one-dimensional flow governing equations coupled with finite-rate chemistry and thermophysical properties tables in the numerical domains of the combustor and nozzle components. The scramjet design space is generated by varying parameters which are flight Mach number and altitude, intake truncation angle, intake exit Mach number and equivalence ratio. Performance outputs of total pressure recovery factor, compression ratio, captured air mass flow rate, intake startability index, thrust, specific impulse, fuel consumption and overall efficiency are computed for each design scenario. The generated database is visualized via performance maps and analyzed in terms of propulsive characteristics. A feature importance study is also conducted to quantify the effects of design parameters on the propulsive performance. • Reduced-order modeling of propulsion systems decreases the complexity of analysis. • Scramjet design space consists of thermal choking, non-ignition, and propulsive zones. • Uninstalled thrust strongly depends on combustor inlet and cruise Mach numbers. • Total pressure recovery is the dominant factor dictating the intake efficiency. • Low speed flow at burner inlet yields autoignition but may cause thermal choking. [ABSTRACT FROM AUTHOR]

The importance of the fuel injection configuration on the propulsion efficiency of high-speed vehicles is apparent. In this article, the use of an annular extruded 4-lobe nozzle for the injection of fuel jet in a supersonic combustor of a scramjet engine in the existence of a shock generator is examined. The main aim of this study is to obtain the efficient jet arrangement for efficient fuel mixing inside the engine of hypersonic vehicles. A numerical approach is used to model the supersonic air stream and cross-jet flow with the SST turbulence model. The role of nozzle altitude and internal air jet on the fuel mixing of the hydrogen within the high-speed domain are disclosed. The importance of the horseshoe vortex and counter-rotating vortex on the fuel distribution is also presented. Our results show that the usage of a coaxial jet instead of an annular jet would increase fuel mixing by more than 40% in the combustion chamber. [ABSTRACT FROM AUTHOR]

Rudder shaft, connecting the air fin to the fuselage of hypersonic vehicles through the mounting gap, is subject to intense aerodynamic heating, necessitating thermal protection techniques. Film cooling is a promising technology to mitigate this heating, but its thermal protection performance on the rudder shaft, as well as aerothermal mechanisms behind major design parameters, is still unexploited. This paper conducts numerical simulations to evaluate thermal protection performance of discrete film cooling applied to the rudder shaft of an air fin in hypersonic flow. Three discrete cooling holes are arranged on the flat plate upstream of the rudder shaft. Parametric studies are then executed to investigate the effect of cooling hole locations and coolant injection angles on thermal protection performance of the rudder shaft. k-ω SST model is used to solve the Reynolds-Averaged Navier-Stokes equations, and perfect gas without chemical reaction is adopted. Total pressure of the freestream is set as 0.6 MPa, while that of the coolant is fixed at 12000 Pa. The freestream has a Mach number of 6 and unit Reynolds number is 7.65 × 106/m. It is found that discrete film cooling can provide excellent thermal protection performance on the rudder shaft by consuming a coolant of 1.41 × 10−4 kg/s. For vertically-injected coolant, thermal protection performance deteriorates as the cooling holes are placed further upstream of the rudder shaft and the distribution of coolant alters qualitatively from a concentrated pattern to a segregated one due to the lower local blowing ratio. If the location of cooling holes is fixed, as the injection angle of coolant decreases, the impingement point of the cooling jet changes from the bottom of air fin to the frontal surface of rudder shaft, and more coolant is attached to the flat plate at the gap bottom. [ABSTRACT FROM AUTHOR]

Curvature plays a crucial role in evolving supersonic cooling film flow-field structures. Flow-field structural images were captured using nanotracer-based planar laser scattering,, and wall pressure values were obtained using experimentally validated numerical simulations. A supersonic cooling film is tangentially injected at the Mach number of Maj = 2.3 into a laminar boundary layer at a mainstream of Ma = 6. The supersonic cooling film inhibits mixing-layer instability on the convex curved wall (CV) and promotes it on the concave curved wall (CC). After increasing the total incoming pressure, the reduction ratio of static pressure (RSP) between the supersonic cooling film and the mainstream flow causes a delay in the position of the mixing-layer instability, smaller-scale vortex structures, and decreased flow velocity of the typical vortex structures on the CC and CV. The wall pressure increases for the CV and decreases for the CC, indicating that the supersonic cooling film suppresses the changes in wall pressure due to curvature. The supersonic cooling film suppresses the decrease in the impulses for bulk dilatation (Ip) due to convex curvature and the increase in Ip due to concave curvature. The growth rate of Ip on the CC increases from −15% to −8% and decreases on the CV from 31% to 12% in the bending impulse (IΦ) range of |IΦ| = 1.337–3.624 for a total inlet pressure of 0.5 MPa. Increasing the RSP could control the Ip values on curved surfaces more effectively. The results of this study can be applied to cooling the infrared optics window on hypersonic vehicles. [ABSTRACT FROM AUTHOR]

The paper addresses enhanced braking efficiency in electromechanically driven vehicle wheels on pneumatic tires through the robust Simple Adaptive Control method. The sliding friction is modeled as dry, semi-dry, and fluid friction, accounting for uneven motion, self-oscillations, and elastic properties of the pneumatic tire. A robust simple adaptive control method is proposed, utilizing signal-parametric adaptation algorithms with implicit reference models. The control objective aims to ensure the convergence of a quadratic performance index within a finite time. The control algorithm combines parametric and signal components, with simulations demonstrating its efficiency. [ABSTRACT FROM AUTHOR]

In the progress toward net zero, two propulsion systems are attracting increasing interest, one based on energy stored in batteries, and the other based on energy stored in hydrogen. In between the applications where hydrogen energy storage has more advantages, there is hypersonic propulsion, which is the object of this perspective. Hypersonic vehicles are aircraft or missiles able to travel at speeds above Mach 5. There are four main types of hypersonic vehicles, spaceplanes, glide missiles, cruise missiles, and aircraft. While glide missiles generally do not need propulsion, cruise missiles, and aircraft are designed to sustain hypersonic speeds for extended periods. They use advanced propulsion systems, which include scramjets, relying on the forward motion of the vehicle to compress incoming air for combustion, and featuring supersonic combustion. While scramjets work well at hypersonic speeds, they must be integrated with other propulsion systems, such as ramjets, rockets, and turbojets, to move back and forth from zero speed. Spaceplanes have been designed so far by using only rockets or integrating rockets and turbojets for better operation in the atmosphere. Spaceplanes' operation in the atmosphere could also be further improved by using ram/scramjets. Point-to-point hypersonic travel for civil applications, which is the focus of this work, requires significant advantages in jet propulsion systems that have to work over different regimes, from subsonic to transonic to supersonic to hypersonic. The work reports on the challenges of hydrogen hypersonic combined cycle propulsion. • Hydrogen scramjet propulsion is necessary for hypersonic aircraft working above Mach 5. • Standalone scramjet propulsion has been successfully demonstrated in very few cases. • Propulsion systems that integrate scramjets with turbojet and ramjet are under development. • Combined cycle propulsion system may deliver faster travel with reduced economic and environmental costs. [ABSTRACT FROM AUTHOR]

Multi-objective optimization of a hypersonic airbreathing engine (scramjet technology) was carried out with the aim of maximizing thrust and minimizing drag while satisfying a series of design constraints, such as avoiding unstart (blockage of supersonic flow within the combustion chamber) by ensuring that the pressure ratio across the shock waves remains below the adverse pressure gradient given by the Korkegi limit, geometry correction to achieve shock on-lip condition, and temperature and pressure requirements at the inlet exit. Using the relations presented in the literature, pressure and viscous drag are estimated analytically. The analytical approach is verified against computational fluid dynamics data from Ansys Fluent to solve two-dimensional compressible Reynolds-averaged Navier–Stokes flow equations, with transition shear stress transport as the turbulence closure model. Comparing the total drag and the flow properties at the combustion chamber entrance shows the model's feasibility for the optimization approach. Three different approaches were conducted to formulate the multi-objective function to determine the one that can find the highest number of geometries satisfying the Korkegi limit with the highest net thrust. The best approach was the multi-objective function formulated with the uninstalled thrust, total pressure recovery, and pressure drag, concentrating the search in the region with greater uninstalled thrust and lower drag and nearly doubling the value of net thrust compared to the first formulation, which uses the uninstalled thrust, pressure drag, and viscous drag. [ABSTRACT FROM AUTHOR]

The issue of hypersonic boundary layer transition prediction is a critical aerodynamic concern that must be addressed during the aerodynamic design process of high-speed vehicles. In this context, we propose an advanced mesoscopic method that couples the gas kinetic scheme (GKS) with the Langtry–Menter transition model, including its three high-speed modification methods, tailored for accurate predictions of high-speed transition flows. The new method incorporates the turbulent kinetic energy term into the Maxwellian velocity distribution function, and it couples the effects of high-speed modifications on turbulent kinetic energy within the computational framework of the GKS solver. This integration elevates both the transition model and its high-speed enhancements to the mesoscopic level, enhancing the method's predictive capability. The GKS-coupled mesoscopic method is validated through a series of test cases, including supersonic flat plate simulation, multiple hypersonic cone cases, the Hypersonic International Flight Research Experimentation (HIFiRE)-1 flight test, and the HIFiRE-5 case. The computational results obtained from these cases exhibit favorable agreement with experimental data. In comparison with the conventional Godunov method, the new approach encompasses a broader range of physical mechanisms, yielding computational results that closely align with the true physical phenomena and marking a notable elevation in computational fidelity and accuracy. This innovative method potentially satisfies the compelling demand for developing a precise and rapid method for predicting hypersonic boundary layer transition, which can be readily used in engineering applications. [ABSTRACT FROM AUTHOR]

Shock drag and aeroheating are major problems faced by hypersonic vehicles. This study proposes a novel three-dimensional model combining spike-aerodisk with a bleed air channel featuring multiple lateral nozzles for drag and aeroheating reduction. The high-temperature and high-pressure air at the stagnation point of the aerodisk is introduced into the channel and ejected laterally through multiple nozzles to modify the flow field structure and reduce drag and aeroheating. In contrast to previous research employing a two-dimensional axisymmetric assumption, our innovation considers the structure of the exit nozzles and the spacing between nozzles. Due to multiple lateral nozzles, the connection thickness between nozzles can lead to asymmetric flow phenomena. Three-dimensional flow field simulation has been employed with the Reynolds-averaged Navier-Stokes (RANS) equations coupled with the shear stress transport (SST) k-ω turbulence model. The effects of the convergence half-angle of the channel inlet, the divergent angle of the nozzle, and the nozzle position on modifying the characteristics of the flow field are studied. Increasing the convergence half-angle of the channel inlet can further push away the reattachment shock, but when it exceeds 60°, the drag reduction performance degrades. Reducing the divergent angle of the nozzle and increasing the number of nozzles can reduce the drag coefficient and the peak value of the Stanton number. However, when the divergent angle is reduced to 10°, the performance of reducing drag and aeroheating is weakened. In the range of research parameters, when the convergence half-angle is 60°, the divergent angle is 15°, and the nozzles are located in the middle of the spike, compared with the plain spike-aerodisk model, the drag coefficient and peak Stanton number of the combined model are further reduced by 17.8 % and 34.8 %, respectively. Our findings confirm that a multi-nozzle can effectively reduce drag and aeroheating while the channel structure is viable in practice. • A new model of lateral jet without gas source is studied by 3-D numerical simulation. • The gas for the lateral jet is coupled with the outside flow. • A distribution form of multiple lateral nozzles is proposed. • A realistic 3-D simulation is conducted, enhancing its practical applicability. [ABSTRACT FROM AUTHOR]

Complex shock interactions on the lips of hypersonic vehicles typically induce severe aerodynamic thermal loads. Considering a V-shaped blunt leading edge as the simplified model of the lip, the design, optimization, and performance of secondary recirculation jets for reducing aerodynamic thermal loads are investigated. The secondary recirculation jet scheme is crucial in reshaping the shock interaction structures, allowing the high-pressure gas downstream of the main shock structures to be transported back into the upstream flow field. Numerical results show that setting the bleed hole of the secondary recirculation jet scheme upstream of the shock wave/boundary layer interaction region yields the greatest reduction in heat flux. The combined effect of the oblique shock upstream of the blow hole and the bow shock at the windward edge of the bleed hole weakens the shock wave/boundary layer interaction structure, thereby reducing the associated heat flux peak. Prior to optimization, the scheme reduces the heat flux peak by approximately 25%. Once the critical design parameters have been optimized and a stagnation bulge has been introduced, the scheme achieves excellent heat flux peak reduction ability of approximately 43%. Simulations under a wide range of freestream conditions show that the optimal scheme has strong applicability to nonzero attack and sideslip angles and is suitable for hypersonic flows below Mach 9.0. In general, the secondary recirculation jet scheme designed in this work exhibits excellent ability to reduce the peak and average values of the heat flux. • A new passive heat flux reduction scheme suitable for the VSBLE model is designed and optimized. • The bleed hole of the secondary recirculation jet scheme is preferably set upstream of the SWBLI region. • The new scheme has satisfactory heat flux reduction ability of the peak and average values after optimizing. • The optimal scheme is suitable for flows with attack angle, sideslip angle, and wide Mach number. [ABSTRACT FROM AUTHOR]

The extensive involvement of blowing in ablation and transpiration cooling directly influences the hypersonic boundary layer (HBL) transition. This work investigates the coupled effect of stabilization through micro-blowing and the corresponding porous blowing medium on a Mach 6 HBL flow. Linear stability theory (LST) and the eN method are used to interpret the characteristics of stability, and direct numerical simulations are used to resolve the detailed flow field among the porous microstructures to verify the predictions of LST. The sole effect of micro-blowing is first investigated by emphasizing the locations of the blowing strip. The results show that micro-blowing alone can significantly excite the first mode but stabilize the second mode. In general, the HBL becomes unstable if the blowing strip is installed upstream of the synchronization points of the dominant disturbances, and otherwise becomes stable. In the context of the blowing medium, a porous coating can suppress high-frequency disturbances and help stabilize the HBL if the blowing strip is placed upstream of the synchronization points of the dominant disturbances. On the contrary, the coating can prematurely excite lower-frequency disturbances and degrade the overall stabilization when the strip is located in the dominant region of the second mode. [ABSTRACT FROM AUTHOR]

A comprehensive numerical study is performed to investigate gas flows inside the inlet of the atmosphere-breathing electric propulsion system using the direct simulation Monte Carlo method. The effects of inlet geometries and gas-surface interaction (GSI) accommodation coefficients on gas pressure and flow rate are analyzed in depth, and a comprehensive assessment of the compression and collection performances of inlets is conducted. Inlet geometries have noticeable effects on gas pressures inside inlets, and the gas pressures of the two concave inlets are the highest. Among the four inlets considered here, the rounded concave inlet performs the best in terms of compression and collection performances. Relative to the simple slope inlet, the rounded concave inlet achieves a rate of increase of over 120% in gas pressure and more than 110% in gas flow rate. GSI accommodation coefficients play a crucial role in both the pressure and mass flux inside the inlet, and the drop of the GSI accommodation coefficient from 1 to 0.2 brings about a considerable increase in the pressure and mass flux, achieving an increase of a factor of 8 and 5, respectively. Therefore, smoothing the inlet surface is an effective means of improving the compression and collection performances of inlets. [ABSTRACT FROM AUTHOR]

We present a discussion on theoretical, experimental, and computational research studies on Görtler instability and the related transition to turbulence occurring in compressible boundary layers over concave surfaces. We first examine the theoretical results on primary and secondary instabilities, emphasizing the role of receptivity, the mechanism by which external agents, such as freestream fluctuations or wall roughness, act on a boundary layer to trigger Görtler vortices. We review experimental findings obtained from measurements in supersonic and hypersonic wind tunnels and discuss studies employing numerical methods, focusing on the direct numerical simulation approach. The research in these two last sections is surveyed according to the geometrical configuration, from simple concave walls to more complex surfaces of hypersonic vehicles. The experimental investigations have been successful in the visualizations of Görtler vortices, in the measurement of the wall-heat transfer in the transitional region, and in the computation of the Görtler-vortex growth rates, although detailed boundary-layer velocity measurements are still missing. Direct numerical simulations have confirmed instability results emerging from stability theories and revealed nonlinear interactions between Görtler vortices and other disturbances. The established initial-boundary-value receptivity theory can certainly benefit from more advanced experimental measurements, and receptivity results should be used in combination with direct numerical simulations. A major conclusion of our review is therefore that the understanding of Görtler vortices should be pursued by a combined methodology including theoretical analysis based on the receptivity formalism, direct numerical simulation, and experiments. Highly desirable outcomes of such endeavor are the prediction of the location and extension of the transition region, and a model for the transition process. We finally highlight further prospects and challenges on fundamental and applied research on Görtler instability and transition in compressible flows. [ABSTRACT FROM AUTHOR]

This paper focuses on the numerical investigation of the transverse hydrogen injection technique in the supersonic airflow on the combustor performance of a scramjet technological demonstrator developed at Universidade Federal do Rio Grande do Norte (UFRN). The UFRN's scramjet vehicle was designed to be fully operational at a hypersonic velocity of 2050 m/s (Mach number 6.8), at 30 km of altitude, which provided a temperature of 1156.95 K and supersonic velocity of 1414.39 m/s (Mach number 2.1), at the combustion chamber entrance. Hydrogen was transversally injected into supersonic airflow at a sonic speed of 1318.46 m/s. A set of 2D-RANS simulations of the reactive flow was performed in combination with a simplified chemical kinetics approach based on the finite-rate chemistry model, defined by a one-step global reaction based on Arrhenius expression to combustion modeling. The turbulence was modeled by the k-kl-ω transition model. Different transverse fuel injections with different dimensions and positioning of injectors were evaluated keeping the same inlet conditions: single-injection versus dual-injection modes. A detailed visualization of several flow phenomena, such as the formation of shock waves, and the performance combustion parameters on the combustor surface were discussed. The analyses indicated hydrogen combustion with a predominantly supersonic flow regime. The transversal injection caused the detachment of the boundary layer as indicated by the recirculation flow structures in the neighboring region next to the injectors. Flow analyses indicate that the purely oblique shock trains were established tending to redirect the airflow over the mixing layer and accelerate the hydrogen burning. Nevertheless, the results showed that this was a local effect since the downstream both mixing and combustion efficiency increased. The dual-injection mode proved to be a promising injection method due to the higher global fuel penetration, elevated from 27,9% to 33,8%, and higher combustion efficiency, increased from 10,33% to 21,99%, with reduced addition of total pressure loss, concerning the single-injection case. Additionally, the growth behavior of performance curves showed that an extension of the combustor downstream length can be indicated to maximize combustor performance. • Two modes of hydrogen transverse injection are studied: single versus dual-injection. • A detailed analysis of the supersonic combustion performance is discussed. • The incidence of shockwaves over mixing layer accelerates the hydrogen burning. • Dual-injection mode provided higher global fuel penetration and combustion efficiency. • Combustor length extension is indicated to maximizing the combustion performance. [ABSTRACT FROM AUTHOR]

This paper presents a novel unscented Kalman filter (UKF) implementation with adaptive covariance matrices (ACMs), to accurately estimate the longitudinal and lateral components of vehicle velocity, and thus the sideslip angle, tire slip angles, and tire slip ratios, also in extreme driving conditions, including tyre–road friction variations. The adaptation strategies are implemented on both the process noise and measurement noise covariances. The resulting UKF ACM is compared against a well-tuned baseline UKF with fixed covariances. Experimental test results in high tyre–road friction conditions show the good performance of both filters, with only a very marginal benefit of the ACM version. However, the simulated extreme tests in variable and low-friction conditions highlight the superior performance and robustness provided by the adaptation mechanism. [ABSTRACT FROM AUTHOR]

This paper investigates the impact of transpiration on a hypersonic flow over a cylinder, considering different degrees of rarefaction. The study analyzes the interaction between freestream argon gas flow at Mach 5 and transpiring argon gas at the fluid–solid interface at a velocity of 10 m/s. Freestream Knudsen numbers considered are 0.002, 0.01, 0.05, and 0.25, spanning from a continuum to rarefied regime. Flow simulations utilize the open-source direct simulation Monte Carlo solver, Stochastic PArallel Rarefied-gas Time-accurate Analyzer. The influence of transpiration on flow and surface properties is examined by comparing non-transpiration and transpiration cases. At all regimes, transpiration increases the normal shock stand-off distance, while a comparison of flow properties along the stagnation line reveals a reduction in the velocity and an increase in the post-shock temperature with transpiration. Surface heat flux comparison indicates that transpiring gas reduces heat flux on the cylinder's upstream-facing front surface at all Knudsen numbers. However, at Kn ∞ = 0.25, a shift occurs, and surface heat flux starts increasing locally from the top/bottom point on the cylinder surface through the rear face of the cylinder. Furthermore, a test for the validity of the continuum-based blowing correction correlation function reveals the failure of the empirical model, even in the continuum regime at Kn ∞ = 0.002, casting doubt on its applicability to vehicles with curvilinear blunt-body shapes. Additionally, a sensitivity analysis demonstrates that transpiring gas with a number density an order of magnitude higher than the freestream reduces stagnation peak heat flux by nearly 30%, while transpiring gas with a temperature two times higher than the freestream shows a ∼13% reduction. [ABSTRACT FROM AUTHOR]

Thermal protection is a crucial issue for a long-flying hypersonic aircraft. In this work, a composite cooling method is proposed and implemented on a hypersonic blunt cone. The composite cooling method combines impinging and convective cooling at the cone head with transpiration cooling employed downstream to produce an overall cooling effect of the cone. Using computational fluid dynamics, the influences of the cooling gas and different attack angles on the cooling effect are examined. The results indicate that the combined cooling method can effectively reduce the overall outer wall temperature of the cone. In particular, with the combined cooling method, an increase in mass flow rate of the cooling gas has been observed in numerical tests, which results in a decrease in the head temperature. The maximum temperature decrease can reach 77.0% on the wall when the mass flow rate of the cooling stream is 1.1 kg/m2 s. The cooling performance on the leeward side can be better than that on the windward side for a fixed cooling gas mass flow rate. This gap in cooling performance between the two sides can be further amplified by a larger attack angle. Quantitatively, when using the combined cooling method, the temperature difference between the windward and leeward sides increases from 106 to 270 K when the attack angle increases from 4° to 8°. The numerical results in this study could provide theoretical and statistical guidance for the design of novel active thermal protection methods for hypersonic aircraft. [ABSTRACT FROM AUTHOR]

Hypersonic stage separation is a significant process, probably involving complex aerodynamic interaction, which determines the survival of two-stage-to-orbit (TSTO) vehicles. The longitudinal stage separation (LSS) scheme is proposed to improve the safety of the parallel-staged TSTO vehicle, where a small interstage gap may result in weak interference. Therefore, an experimental and numerical study of LSS for the parallel-staged TSTO vehicle at Mach 7 with different angles of attack (AoA) is carried out. The dynamic interaction, including variations in the shock structure, wall pressure distribution, and unsteady aerodynamics, is investigated by testing and numerical simulation. The LSS experiments for the TSTO vehicle were performed using a high-speed pneumatic ejection launch system in the JF-12 shock tunnel, and the method was developed using high-speed visualization and image processing techniques to capture the separating trajectory. The numerical simulations were carried out using the overset grid method and solving the Navier–Stokes equations coupled with the rigid body dynamics equations to obtain the laminar flows over the TSTO vehicle during LSS. The qualitative and quantitative comparison of the test and numerical results showed good agreement in terms of aerodynamic performance, flowfield pattern, wall pressure, and separation trajectory. They show that the small interstage gap of the LSS leads to weak type I and VI shock–shock interactions, with short-duration weak shock reflection at a higher AoA. Furthermore, no shock reflection or interstage gap is observed at lower AoA. Moreover, no stage recontact is observed, and the safety and feasibility of LSS for parallel-staged TSTO vehicles are demonstrated. [ABSTRACT FROM AUTHOR]