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As for blunt-body hypersonic vehicles, the extreme drag and aero-heating pose great challenges to the conceptual design. Different from the conventional configurations, a blunt body slotted with a channel is proposed in this paper. The channel is further combined with a jet on the shoulder to reduce drag and aero-heating. Simultaneously, based on the position of the separation point and vortex edge, a modified effective body approach is developed to accurately investigate the drag reduction mechanism. Specifically, the more slender the modified effective body is, the more obvious the drag reduction is. The simulation results show that the spike-channel-jet can push the shear layer on the spike away from the wall. In this way, the position of shear reattachment point moves downstream, which can enlarge the recirculation zone and weaken the shock-shock interaction. Additionally, the reattachment shock is pushed far away from the blunt body, which is also weakened. For drag, net power is introduced to evaluate the practical value of spike-channel-jet for drag reduction. Compared with a simple spike (spike length L / D = 1) of the same size, the spike-channel-jet can reduce the total drag by 24.68 %, the peak pressure coefficient of the blunt body by 38.54 %, the peak temperature of the blunt body by 65.75 %, and the peak temperature of the spike by 10.63 %. The maximum net power can reach 13.30 kW. Considering the fluid-thermal interaction, the findings reveal the influence of the spike length and jet mass flow rate on aerodynamic performance, by modified effective body further analyze the drag reduction mechanism with the jet, and validate the different reasons for the peak temperature and pressure of the blunt body. • Drag and aeroheating reduction mechanism of the spike-channel-jet concept under hypersonic conditions is investigated by the numerical simulation. • The influence factors are analyzed and compared comprehensively. • Combined spike-channel-jet concept is a promising drag and aeroheating reduction technique. [ABSTRACT FROM AUTHOR]

Currently, aerodynamic environment prediction research into scramjet-propelled vehicles characterized by NACA0012 under hypersonic conditions is relatively sparse. Two-dimensional external flow field models are established, and then through validation tests, we perform a systematic investigation between simulation parameters and prediction accuracy, and an effective aerodynamic environment prediction simulation scheme under hypersonic conditions is proposed. Unlike under incompressible conditions, the maximum accuracy decline could be attributed to the inappropriate choice of the sharp trailing edge modeling method, but the definition formula is still preferred. In particular, for the two modeling data point sources, Airfoil tools and NACA4, the numerical performance of the latter is better than the former, and the calculation accuracy negatively correlates with the number of data points offered by both of them. Moreover, for the mesh cells near the shock, the cell Reynolds number and aspect ratio values should be no smaller than 16 and not exceed 380, respectively, and the recommended values for the far field distance, the turbulence model and flux type are 16L, Spalart-Allmaras, and ROE flux type. Under hypersonic conditions, the aerodynamic environment characterized by NACA0012 predicts a maximum temperature of approximately 1856.85 °C, with an average temperature change rate of 77 °C/s. Meanwhile, the top sound pressure level and the vibration acceleration could reach up to 145 dB and 182 g, respectively. [ABSTRACT FROM AUTHOR]

Pulse opposing plasma synthetic jets (OPSJs) have potential for reducing the drag coefficient of hypersonic airfoils. This paper numerically investigated the effects and flow control mechanisms of a continuous nanosecond pulse OPSJ on the aerodynamic performance of a hypersonic airfoil. The results indicate that a continuous-pulse OPSJ reduces the airfoil drag coefficient by generating diffracted waves to shift the leading-edge shock wave. However, the flow control effectiveness of a pulse OPSJ decreases with the increase of the pulse sequence. The time-averaged lift-to-drag ratio of the airfoil decreased rapidly from 1.11 for the first pulse sequence to 0.94 for the third pulse sequence, followed by a gradual decrease. The thermal accumulation within the cavity triggers a blocking effect, impeding the heat exchange between the cavity and external fluid. This is the primary factor leading to the decreasing flow control effectiveness of continuous-pulse OPSJs. Furthermore, the results suggest that enhancing the mixing effect within the cavity can improve heat transfer between the cavity and the external flow, suppress thermal accumulation, and improve the flow control capability of continuous-pulse OPSJs in hypersonic flows. [ABSTRACT FROM AUTHOR]

Copyright of Journal of Ballistics / Dandao Xuebao is the property of Journal of Ballistics Editorial Department and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

This paper details the development and verification tests of an open-source code named hy2LPTFoam intended for solving high-speed non-equilibrium gas-particle flows in OpenFOAM. The solver, based on hy2Foam for high-speed non-equilibrium gas flow, integrates multiple particle force models, heat transfer models, the diffusion-based smoothing method, and the MPPIC method. The verification tests incorporate interactions between shock waves and particle curtains with varying particle volume fractions, a JPL nozzle generating a two-phase gas–particle flow, and a Mars entry body with two particle inflow mass fractions. The tests yield good physical agreement with numerical and experimental data from the literature. [ABSTRACT FROM AUTHOR]

Predicting the gas-surface interactions of solid carbon is necessary for the design of many engineering systems that employ graphite. Experimental determination of the reaction rates improves the fidelity of those predictions. Here, we studied oxidation and nitridation of graphite by thermal and nonthermal plasma assisted processes. Experiments were conducted at a pressure of 2 kPa, higher than previous flow reactor experiments of this kind and closer to the conditions experienced in engineering applications. At these higher pressures, the limitations of mass transport and the interference between oxygen and nitrogen species become important. Reaction rates were determined from mass loss, reaction products were identified with mass spectrometry, and surface roughening was characterized by electron microscopy. [ABSTRACT FROM AUTHOR]

With the rapid development and application in practical battlefields of hypersonic weapons, the demand for hypersonic defense technologies has become increasingly urgent in various countries. Due to the different target characteristics of hypersonic weapons from traditional missiles, existing missile defense systems are difficult to achieve good interception effects, and related equipment and technologies need to be redeveloped. In order to sort out the future demand directions of hypersonic defense technologies, this paper summarized the current status of hypersonic defense capabilities in countries such as the United States, Russia, and Japan from the aspects of early warning detection, interception and strike and so on, and analyzed relevant key technologies. It is proposed that future hypersonic defense should focus on increasing interception distance, reducing combat costs, and dealing with system attack.

Accurate heat flux prediction is a key problem in the numerical simulation of hypersonic flows. When the singularity axis appears in the grid, it further increases the difficulty of accurate flow simulation and heat flux calculation. Aiming at the problems of instability of flow structure and large errors of heat flux prediction in the finite difference calculation of singularity axis problems, a hybrid algorithm for solving the singularity axis problem with high precision difference schemes is proposed by combining the flux reconstruction and conservation reconstruction. The algorithm is based on conservative variables reconstruction near the singularity axis, which can reflect the influence of the flow field distribution near the singularity axis. The algorithm can effectively solve the singularity axis problem in numerical simulation. Applying the hybrid algorithm with WENO3-JS and WENO3-PRM scheme, the numerical simulations of hypersonic flow around sphere, HB-2, sphere cone and Apollo are carried out. The results show that the new method can effectively solve the problems of flow instability with singularity axis, and can achieve accurate heat flux prediction. [ABSTRACT FROM AUTHOR]

Copyright of Aero Weaponry is the property of Aero Weaponry Editorial Office and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

In order to determine the installation position of the front fin on the example rocket forebody, an optimized method based on a comprehensive evaluation indicator and variable step search is presented. The comprehensive indicator consists of four weight coefficients, two lateral aerodynamic forces and two aerodynamic moments. The computational fluid dynamics method based on a shear stress transport turbulence model is established to analyze the flow field characteristics of the forebody. The results indicate that under equal weight coefficients, the presented search algorithm can provide an optimized solution for the front fin to achieve the minimum value of the comprehensive evaluation indicator. When the range of the current wing movement changes or the weight coefficient distribution changes, this search algorithm can still provide the optimal solution and some feasible solutions. Under the given conditions, there is a difference between the optimal solution of the aerodynamic force priority and that of the aerodynamic moment priority. For the case of the aerodynamic moment priority, the mean level of the pressure coefficient corresponding to the optimal solution on the given observation plane is low. The presented method is effective in learning the appropriate installation position of the rocket's front fins. [ABSTRACT FROM AUTHOR]

This paper studies the local receptivity of the Mack-mode instability to localized unsteady blowing and suction (UBS) in a chemical non-equilibrium (CNE) hypersonic boundary layer. The five-species CNE model is employed, and the receptivity efficiency is formulated by use of the residual theorem. Compared with the results for the calorically perfect gas (CPG) model, we find that the real-gas effect enhances the receptivity efficiency remarkably in the majority of the second-mode frequency band, and the enhancement is mainly attributed to the modification of the base flow due to the CNE effect, which is akin to the cold-wall effect in hypersonic boundary layers. Combined with the destabilizing role of the CNE effect on the Mack second mode, it is concluded that the CNE effect would lead to a greater linearly accumulated perturbation amplitude, implying premature of transition to turbulence in a high-enthalpy hypersonic boundary layer subject to localized perturbations. [ABSTRACT FROM AUTHOR]

High-speed global surface temperature distributions and heat flux measurements on the Initial Concept 3.X vehicle (IC3X) model were investigated at the UTSA Mach 7 wind tunnel, examining angles of attack of 0° and 5° at a freestream unit Reynolds number (Re) ~7 × 106 m−1. A ruthenium-based, fast-responding, temperature-sensitive paint (fast-TSP) prepared in-house was applied to a 7.1% scale model of the vehicle. Static calibration was performed to convert the intensity measurements into surface temperature values. The surface temperatures and derived heat flux fields conformed to the predicted trends, which was corroborated by Schlieren flow visualization. Notably, the average surface temperature variation was identified to range from 6 to 34 K at a 0° angle of attack and from 11 to 44 K at a 5° angle of attack, with the most pronounced gradient detected at the stagnation point. Additional measurements provided a detailed thermal assessment of the model, including estimations of the stagnation point heat flux, the convective heat transfer coefficient, and the modified Stanton number. Statistical and time series analyses of the data collected revealed the absence of prevailing unsteady phenomena, suggesting that the tested design geometry is well suited for hypersonic flight applications. These experimental outcomes not only shed light on the aerothermodynamics experienced during high-speed flight but also underscore the effectiveness of fast-TSP in capturing both quantitative and qualitative thermal data. [ABSTRACT FROM AUTHOR]

In order to study the variation in the lift performance as a function of the angle of attack of a hypersonic gliding vehicle under supersonic (hypersonic) conditions, the cuspidal waverider was taken as the object of this study, and the variation in the lift performance depending on the angle of attack was simulated for its design condition (M = 3.86 and H = 25 km) and hypersonic incoming flow condition (M = 8 and H = 25 km). It was also compared with the delta-wing Model 1 with the same leading-edge swept angle and Model 2 with the same spread length, respectively. The obtained results show that the zero-lift angle of attack and the critical angle of attack of the cuspidal waverider are both greater than those of the Model 1 and Model 2. The critical angle of attack increases with the free-stream Mach number for all three models. The maximum lift coefficient angle of attack on the upper surface of the cuspidal waverider decreases with increase in the Mach number, contrary to Models 1 and 2, and this relates to the degree of expansion of the free-stream flow conditions, the model layout, and the vortex structure formed on the leeward side. [ABSTRACT FROM AUTHOR]

Supersonic and hypersonic flows have gained considerable attention in the aerospace industry in recent years. Flow control is crucial for refining the quality of these high-speed flows and improving the performance and safety of fast aircraft. This paper discusses the distinctive characteristics of supersonic flows compared to low-speed flows, including phenomena such as boundary layer transition, shock waves, and sonic boom. These traits give rise to significant challenges related to drag, noise, and heat. Therefore, a review of several active and passive control strategies is provided, highlighting their significant advancements in flow transitions, reducing drag, minimizing noise, and managing heat. Furthermore, we provide a comprehensive analysis of various research methodologies used in the application of flow control engineering, including wind tunnel testing, flight testing, and computational fluid dynamics (CFD). This work gives an overview of the present state of flow control research and offers insights into potential future advancements.

Abstract Considering the influence of thermal stress and material property variations, this study employs the Navier–Stokes equations and Fourier heat conduction law to establish a semi-implicit time-domain numerical analysis method for hypersonic aerothermal-structural coupling. Study the temporal variation pattern of different regions of the composite material wing under aerodynamic heating. Using the obtained transient temperature field of the wing, the thermal modal of the wing at different time points is calculated using the finite element method. Additionally, it conducts an analysis and discussion on the factors influencing the thermal modal. Composites can be effectively utilized as thermal protection materials for aircraft. During the aerodynamic heating process, the leading edge temperature reaches thermal equilibrium first, followed by the trailing edge, and the belly plate experiences a slower thermal response. Temperature rise significantly affects higher-order modes, with the change in material properties during the early stages of heating being the dominant factor. This leads to a faster decrease in natural frequency. As heat conduction progresses, the influencing factors of thermal stresses gradually increase, and the natural frequency decreases slowly or even rises.

This paper provides an examination of the modern high-speed aircraft industry, including a compilation of recent developments and planned aircraft. The potential for a civil high-speed airliner is examined from a financial and contemporary perspective. Mach 4 is introduced as the economic line, below which high-speed flight is shown as financially viable. Despite this, a preliminary analysis from historic data highlights flight below Mach 2.12 as relatively low risk, with a hypersonic airliner not envisioned until the next century. Several challenges that are less discussed in literature are reviewed here with safety and reliability being the principal among them. This paper hence serves as the first part, which is followed by a comprehensive review of high-speed vehicle accidents.

Considering the influence of thermal stress and material property variations, this study employs the Navier–Stokes equations and Fourier heat conduction law to establish a semi-implicit time-domain numerical analysis method for hypersonic aerothermal-structural coupling. Study the temporal variation pattern of different regions of the composite material wing under aerodynamic heating. Using the obtained transient temperature field of the wing, the thermal modal of the wing at different time points is calculated using the finite element method. Additionally, it conducts an analysis and discussion on the factors influencing the thermal modal. Composites can be effectively utilized as thermal protection materials for aircraft. During the aerodynamic heating process, the leading edge temperature reaches thermal equilibrium first, followed by the trailing edge, and the belly plate experiences a slower thermal response. Temperature rise significantly affects higher-order modes, with the change in material properties during the early stages of heating being the dominant factor. This leads to a faster decrease in natural frequency. As heat conduction progresses, the influencing factors of thermal stresses gradually increase, and the natural frequency decreases slowly or even rises. [ABSTRACT FROM AUTHOR]

The parametric design method is widely utilized in the preliminary design stage for hypersonic vehicles; it ensures the fast iteration of configuration, generation, and optimization. This study proposes a novel parametric method for a wide-range, wing-morphing glide vehicle. The whole configuration, including a waverider fuselage, a rotating wing, a blunt leading edge, rudders, etc., can be easily described using 27 key parameters. In contrast to the typical parametric method, the new method takes internal payloads into consideration during the shape optimization process. That is, the vehicle configuration can be flexibly adjusted depending on the internal payloads; these payloads may be of random amounts and have different shapes. The code for the new parametric design method is developed using the secondary development tools of UG (UG 10.0) commercial software. The lift and drag characteristics over a wide operational range (H = 6–25 km, M = 2.5–8.5, AOA = 0–10°) were numerically investigated, as was the influence of the retracting angle of the morphing wings. It was found that, for the mode of the fully deployed wings, the lift-to-drag ratio (L/D) remained at a high level (≥4.7) over a Mach range of 4.0–8.5 and an AOA range of 4–7°. For the mode of the fully retracted wings, the drag coefficient remained smaller than 0.02 over a Mach range of 4.0–8.5 and an AOA range of 0–5°. A wide L/D of 0.3–4.7 could be achieved by controlling the retracting angle of the wings, thus demonstrating a good potential for flight maneuverability. The flexible change in L/D proved to be a combined result of varying pressure distribution and edge-flow spillage. This will aid in the further optimization of lift/drag characteristics. [ABSTRACT FROM AUTHOR]

Hypersonic stage separation is a significant process for the future two-stage-to-orbit (TSTO) vehicle. Strong and complex interstage aerodynamic interference may result in drastic aerodynamic forces and moments, potentially conflicting with the safe separation. Therefore, the dynamic stability of the vehicle during separation is critical to aerospace safety. In this study, numerical analysis of a hypersonic flow with Ma = 6.7 past a parallel-staged TSTO vehicle during stage separation is performed by laminar flow simulations. The TSTO vehicle consists of a wave-rider and a spaceplane as booster and orbiter, respectively. Considering the different centers of gravity for the orbiter during separation, the longitudinal dynamic stability of the orbiter is analyzed based on the dynamic characteristic of the center of pressure (CoP) in different cases. The dynamic separation behavior, aerodynamic characteristics, and typical flowfield patterns are clarified. Moreover, the derivative of CoP to time is proposed and analyzed in detail, which serves as an indicator to determine the dynamic stability and the safe stage separation. A safety separation judgment criterion is also proposed based on the CoP dynamic characteristics for the TSTO vehicle, and the mechanism of CoP variations associated with the unsteady aerodynamic interference during stage separation is revealed. • The safe margin on the dynamic separation trajectory for the TSTO vehicle is analyzed. • The typical flowfield patterns with unsteady interaction for TSTO stage separation are clarified. • The dynamic stability analysis methods for TSTO stage separation based on CoP are performed. • The general safe separation criteria based on the stability analysis method for TSTO are analyzed. • The unsteady interaction mechanism associated with dynamic stability analysis is emphasized. [ABSTRACT FROM AUTHOR]

Near-space hypersonic vehicles could experience both continuous and rarefied flow regimes during the flight through atmosphere. The rarefaction effects on hypersonic boundary-layer stability are studied based on the Navier-Stokes (NS) equations and an improved NS model. The conventional linear stability theory (LST) is extended for rarefied shear flows by adopting slip boundary conditions and nonlinear transport relations. A flow at Mach 10 over a flat plate at an altitude of 55 km is the main case of analysis. The separate and combined effects of rarefaction (including surface slip and shear nonequilibrium) on stability by influencing the base flow and stability equation are clarified. The results show that for the base flow, rarefaction effects cause the boundary layer to become thinner and the generalized inflection point to move towards the wall. For stability, rarefaction effects have a stabilizing effect on the second-mode instability by influencing the base flow while a destabilizing effect by modifying the stability equation. The combined effects of rarefaction suppress the second-mode instability for different Mach number cases. However, for the first-mode instability, rarefaction effects play a destabilizing role. These results shed light on the hypersonic boundary-layer stability in the near-continuum regime from a macroscopic view. [ABSTRACT FROM AUTHOR]

Hypersonic flight is of extreme interest today because of a new vehicle concept designed to fly at very high Mach numbers. In hypersonic flows, strong shocks are generated with a thicker boundary layer around the surface of flights. In hypersonic flows, nose bluntness for the leading edges of the aerofoil is extremely useful for maximum lift to drag ratio to reduce aerodynamic heating. In hypersonic flows, space between the blunt body nose and the shock generated is mixed subsonic and supersonic flow through the sonic line. The only CFD method available today is the time marching method, as the unsteady equations of motion are hyperbolic with respect to time regardless of whether the flow is locally subsonic or supersonic. ANSYS ICEM CFD is used for meshing two and three-dimensional models as it is having a powerful blocking method. Fluent was used as a solver and postprocessor. A computational study of flow around a power law body was conducted for a Mach number of 8.2, pressure = 951.5 N/m2, temperature = 89.3 K and velocity = 1553 m/s. The Mach number contours obtained were close to experimental values. Coefficient of Drag (CD), Coefficient of Lift (CL) and Coefficient of Moment (CM) obtained were matching to the experimental results. [ABSTRACT FROM AUTHOR]

Focusing on the plasma wakes of hypersonic targets (PWHT), this work provides a method for the inversion of electron densities in the PWHT by integrating electromagnetic radiation theory and three-wave coupling theory (TWCT). Utilizing the Computational Fluid Dynamics (CFD) method, taking a blunt cone as an example, the spatial distribution characteristics of electron densities in the plasma wake are simulated and analyzed in detail. The Zakharov system of equations (ZSE) is applied to the PWHT for the discussion about the electromagnetic radiation generated by plasma parametric instability. Combining the wavenumber spectrum of electromagnetic radiation and the TWCT, an inverse method of the electron densities of plasma wakes is proposed and applied to a blunt cone model. The results indicate that the inversion error rate is effectively held below 1.37 %. Furthermore, reasons for errors are elucidated through the utilization of the wavenumber spectrum structure. The proposed method may contribute to the development of hypersonic wake diagnostic techniques, and offer new insights and directions for wind tunnel experiments, shock tube experiments, as well as research on hypersonic target detection.

This paper details the development and verification tests of an open-source code named hy2LPTFoam intended for solving high-speed non-equilibrium gas-particle flows in OpenFOAM. The solver, based on hy2Foam for high-speed non-equilibrium gas flow, integrates multiple particle force models, heat transfer models, the diffusion-based smoothing method, and the MPPIC method. The verification tests incorporate interactions between shock waves and particle curtains with varying particle volume fractions, a JPL nozzle generating a two-phase gas–particle flow, and a Mars entry body with two particle inflow mass fractions. The tests yield good physical agreement with numerical and experimental data from the literature.

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]

As the flow temperature rises dramatically at high altitudes and hypersonic speeds, the thermodynamic properties of gas change, causing changes in aerodynamic properties of waverider. Therefore, it is of great engineering value to investigate real gas effects in the design of waverider. A method of characteristic (MOC) considering equilibrium gas effect is proposed for the design of cone-derived waverider. Different shapes of waverider are generated based on different half cone angles and different flow capture curves, and the aerodynamic and geometrical characteristics of each waverider under equilibrium gas and perfect gas conditions are estimated using a fast estimation method. The results shows that there are differences in the lift-to-drag ratio, volume ratio and pitch moment coefficient of the waverider designed based on the equilibrium gas model at high Mach number compared with the perfect gas results. The degree and pattern of the influence of equilibrium gas effect on different configurations of waverider are distinct. [ABSTRACT FROM AUTHOR]

An empirical study assessing the suitability of laser speckle patterns for use in experimental mechanics research conducted at hypersonic conditions is presented. Speckle patterns generated via coherent optical scattering from a gold-coated graphite surface inserted into a Mach 10 air stream were collected and analyzed using Computer-Aided Speckle Interferometry using two methodologies: a typical spanwise approach involving increasing temporal span between image pairs, and a sequential approach that utilizes a combination of temporal re-referencing and spatial resampling. Results were characterized for spatial-temporal stability and variance. Our observations demonstrate that optical methods relying on stable and properly formed speckle patterns can be used in a hypersonic regime at various Reynolds numbers and common camera frame rates, but the technique is sensitive to the conditions of the airflow and the processing methodology chosen. [ABSTRACT FROM AUTHOR]

Explorations involving long-endurance and maneuvering flights in the upper atmosphere, as well as research on atmospheric entries of space debris or asteroids, call for a full understanding of hypersonic rarefied flows. The inverse Magnus effect occurs in the hypersonic rarefied flow past a rotating sphere, but the aerodynamic behavior is contrary to the Magnus effect in the continuum flow regime. In this article, a series of such flows are numerically studied using the direct simulation Monte Carlo (DSMC) method. By analyzing the flow fields, as well as the distributions of pressure and shear stress on the sphere, the formation of the inverse Magnus force can be attributed to the tangential momentum transfer between incident gas molecules and the windward surface. The variation laws of aerodynamic parameters with the rotation rate are presented and discussed. [ABSTRACT FROM AUTHOR]

The present work aims to study the effects of shock wave — boundary layer interaction in 2D hypersonic flow at the sharp leading edge of a flat plate and a flat plate followed by a compression ramp. A key parameter in this work is the hypersonic interaction parameter χ , defined as χ ≡ C M ∞ 3 / R e ∞ 1 / 2 . This flow is modeled as a perfect gas and its behavior is studied over various compression ramp configurations, ranging from 0 ° – 6 ° , using a semi-analytic reduced-order model. In this shock-layer analysis we focus on characterizing the viscous boundary layer for an adiabatic flow over an insulated wall, improving on existing semi-analytic methods by obtaining a non-singular solution, and using the Tangent Wedge approximation to match the pressure on the boundary layer edge. Upstream influence resulting from the compression ramp is also investigated with newly developed correlations. [ABSTRACT FROM AUTHOR]

The double-wedge configuration is a typical characteristic shape of the rudder surface of high-speed aircraft. The impact of the shock wave/boundary layer interaction and the shock wave/shock wave interaction resulting from the double wedge on aircraft aerodynamics cannot be ignored. The aerodynamic performance of the aircraft would be seriously affected. Accordingly, to reduce the wave drag, and to relieve the thermal load and pressure load, flow control is required for the shock wave/shock wave interaction and the shock wave/boundary layer interaction induced by the double-wedge configuration. In this paper, double-wedge shock wave/shock wave interaction is controlled by a high-energy surface arc discharge array and observed by high-speed schlieren flow field measurement at Mach 8. The 30-channel discharge array is set on the primary wedge plane, and actuation is generated. Hypersonic V shock wave/shock wave interaction is effectively controlled by the shock wave array induced by the high-energy surface arc discharge array, which makes the shock wave/shock wave interaction structure disappear or intermittent. The potential control mechanism is to reduce strong shock wave interaction by transforming the type of shock wave interaction. Therefore, the ability of plasma array actuation to control complex shock wave/shock wave interaction is verified, which provides a new method for hypersonic shock wave/shock wave interaction control. [ABSTRACT FROM AUTHOR]

Copyright of Journal of Ordnance Equipment Engineering is the property of Chongqing University of Technology and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

In this study, shock stand-off distances for thermally and chemically non-equilibrium flows of nitrogen over wedges are computationally investigated via a hypersonic computational fluid dynamics solver, hyperReactingFoam by spanning a parameter space that consists of ranges of Mach number, 4–10, specific heat ratio, 1.40–1.61 and wedge angles, 60 ∘ –90 ∘ . Then, the space is reduced into the parameters of inverse density ratio across the shock and dimensionless wedge angle which will be used as variables for quadratic functions that represent shock stand-off distances. Besides the functions of shock stand-off distances, detached shock profiles of computationally modeled flows are represented by parabolic equations. The flows are observed to be chemically frozen for Mach number ranges of 4–5 regardless of the specific heat ratio value of the nitrogen mixture. Our results show that the shock stand-off distance decreases as Mach number is increased from 4 to 7, if the wedge angle and free-stream specific heat ratio are kept the same. On the other hand, if Mach number is increased beyond 7, the shock stand-off distance starts to extend due to the dissociation of nitrogen molecules behind the shock wave. At Mach 10, nitrogen completely dissociates over 90 ∘ wedge for all specific heat ratios considered in the present study. Increased leading edge angle of the wedge or specific heat ratio of free-stream yields longer shock stand-off distance. [ABSTRACT FROM AUTHOR]

This research aims at analyzing thermo-chemical properties of the hypersonic high-enthalpy flow in the L2K wind tunnel, situated in Köln at the German Aerospace Center (DLR). In the L2K wind tunnel, Martian atmosphere can be created, and the facility can simulate heat load conditions encountered during atmospheric entry of Martian missions. The focus of this project is the analysis of the non-intrusive experimental technique "Tunable Diode Laser Absorption Spectroscopy" (TDLAS), based on line of sight absorption spectroscopy, and applied to hypersonic flow. A simplified Martian atmosphere (97% CO 2 and 3% N 2) was used. A new interpretation for CO-TDLAS experimental technique applied to hypersonic wind tunnel flow analysis was developed. Numerical simulations with the DLR-TAU non-equilibrium flow solver were used as support of this analysis, and match between simulations and experiments was observed. Flow speed and absorption line's width were measured, and the knowledge of L2K's flow structure was extended. [Display omitted] • TDLAS experimental technique in L2K wind tunnel was comprehensively analyzed. • Recirculation process of the flow's outer layers was confirmed and analyzed. • Match between simulations of spectroscopic data and experiments was observed. [ABSTRACT FROM AUTHOR]

Compression corner shock wave/boundary layer interaction (SWBLI) is a typical shock wave/boundary layer interaction (SWBLI) problem in supersonic/hypersonic flows. In previous studies, the separation flow is usually caused by a single shock wave. However, in the actual aircraft surface configuration, two-stage compression or even multistage compression will produce more complex SWBLI problems. The multi-channel shock structure makes the flow field structure more complicated and also puts forward higher requirements for the flow control scheme. In order to explore a flow control method for the double compression corner shock wave/boundary layer interaction problem, an experimental study is carried out to control the double compression corner shock wave/boundary layer interaction with a high-energy flow pulsed arc discharge array under the condition that the incoming flow velocity Ma 6.0 has both noise flow fields and quiet flow fields. The results show that when UDC = 0.5 kV actuation is applied, the influence range of the hot gas mass flow direction is about 65 mm, which can weaken the shock wave intensity to a certain extent. When UDC = 1 kV actuation is applied, the influence range of the hot gas mass flow direction extends to 85 mm, and the actuation has a significant control effect on the flow field. Through spatio-temporal evolution analysis and spatial gradient threshold processing of high-speed schlieren images of actuated flow fields, the feasibility of controlling the hypersonic double compression corner shock wave/boundary layer interaction by using a high-energy flow pulsed arc discharge array is verified. The control law of a high-energy flow pulsed arc discharge array acting on the double compression corner shock wave/boundary layer interaction is revealed. [ABSTRACT FROM AUTHOR]

Hypersonic inlet is very easy to induce aerothermoelastic problems under the aerodynamic load and aero-thermal action of complex flow. Deeply understanding the aerothermoelasticity mechanism of complex internal flow is of great significance for the detailed design of hypersonic inlet in the future. In this paper, a static/dynamic aerothermoelastic analysis framework was established, and the mechanism of the influence of static/dynamic aerothermoelasticity on the flow field structure and performance of three-dimensional hypersonic inlet was studied in depth. The results of static aerothermoelastic analysis show that the aerothermoelastic deformation obtained by the two-way coupling method is relatively large, and the deformation of the leading edge of the inlet lip is the largest. The structural deformation changes the shock wave structure near the lip edge, enhances the shock wave intensity inside the inlet, increases the length of the separation zone and the temperature of the outer wall, and changes the flow field at the outlet. At the same time, the aerothermoelastic deformation will lead to the increase of mass flow coefficient and pressure rise ratio, and reduce the total pressure recovery coefficient. The results of dynamic aerothermoelastic analysis show that the displacement response of the structure converges when the aerodynamic heating is not taken into account. After considering aerodynamic heating, the structural displacement response presents a limit cycle trend. Aerodynamic heating may change the dynamic response characteristics of the inlet structure. Because the structural frequencies of the intake ports are very close to each other, ″beat″ phenomenon exists in the dynamic response of the structure. The leading edge deformation is large and the amplitude is small, while the trailing edge deformation is small and the amplitude is large. The structure vibration leads to significant dynamic changes in the flow field structure and significant fluctuations in the performance parameters, especially for the pressure rise ratio at the outlet. It is hoped that the research in this paper will deepen the understanding of the aerothermoelasticity in the complex flow structure of the inlet, in order to provide reference for the detailed design of the inlet in the future.

In order to determine the installation position of the front fin on the example rocket forebody, an optimized method based on a comprehensive evaluation indicator and variable step search is presented. The comprehensive indicator consists of four weight coefficients, two lateral aerodynamic forces and two aerodynamic moments. The computational fluid dynamics method based on a shear stress transport turbulence model is established to analyze the flow field characteristics of the forebody. The results indicate that under equal weight coefficients, the presented search algorithm can provide an optimized solution for the front fin to achieve the minimum value of the comprehensive evaluation indicator. When the range of the current wing movement changes or the weight coefficient distribution changes, this search algorithm can still provide the optimal solution and some feasible solutions. Under the given conditions, there is a difference between the optimal solution of the aerodynamic force priority and that of the aerodynamic moment priority. For the case of the aerodynamic moment priority, the mean level of the pressure coefficient corresponding to the optimal solution on the given observation plane is low. The presented method is effective in learning the appropriate installation position of the rocket’s front fins.