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Three-dimensional high-fidelity numerical simulations of a Mach 5 hypersonic ramjet intake are performed combining a high-order and time-accurate large-eddy-simulation model with a sharp-interface ...

This work describes numerical and experimental research in the Indian Hypersonic Shock Tunnel 3 on a blunt cone model and related computation of the facility parameters: nozzle reservoir conditions...

This paper investigates the neural adaptive control problem for air-breathing hypersonic vehicles. For the velocity subsystem, a radial basis function neural network (RBFNN)-based adaptive controller is first designed, which employs the auxiliary variable to compensate for the saturation nonlinearity of the scramjet control command. For the altitude subsystem, an RBFNN-based controller addresses actuator constraints and dynamics using the model predictive control, as well as counteracts uncertainties and disturbances using the neural adaptive mechanism. The effectiveness of the proposed control is verified by simulations.

A quantized fixed-time fault-tolerant attitude control problem for hypersonic reentry vehicles is considered from the perspective of practical engineering. Taking into account the limited time available for recovery from a fault scenario, fixed-time extended state observers are used to simultaneously compensate for the negative effect of unknown time-varying actuator faults, uncertainties, and external disturbances. In contrast to the existing control schemes, which neglect the signal quantization transmission process and limited communication capacity, the hysteresis quantization mechanism is considered. Based on this, a fixed-time fault-tolerant control scheme is developed to ensure the fixed time convergence of all the tracking errors even in cases involving actuator faults. The stability and convergence are proven by performing a theoretical analysis. The simulation results validate the superiority and robustness of the proposed control scheme.

Heat removal capacity of catalytically cracked supercritical n-dodecane as a jet fuel analogue in a cylindrical packed-bed reactor is examined. Fuel cracking is endothermic, and can be used in desi...

Metallic lattice-frame materials, due to excellent heat transfer performance under high heat flux and mechanical property, are regarded as the potential candidates for a vehicle leading edge. This ...

This paper reports an active method for aerodynamic force control of a hypersonic spacecraft via shock wave mitigation with plasma energy deposition. The energy deposition was generated by quasi-direct current discharge at two positions in the streamwise direction on a flat plate upstream a compression surface. Wind tunnel experiments of visualizing a Mach 6 flow over a ramp with high-speed schlieren imaging were performed to verify the method's effect on diluting a shock wave. Meanwhile, the Reynolds averaged Navier-Stokes equations, with a source term of simulating plasma heating due to discharge, were solved to evaluate the method's control abilities. From the experimental results, low-density plasma layers can be formed by quasi-DC discharge, which were continuously weakening the shock wave induced by the ramp. The numerical results predicted flow topology under control well with the experimental schlieren images, with the corresponding force reduction rate as high as 58%, which mainly depends on the amount of energy injection, instead of arrangement of heating zones in this study.

Velocity-slip and temperature-jump effects on sharp leading-edge geometries are studied for three canonical hypersonic flows: the hollow-cylinder-flare, the double-cone, and the double-wedge. Simul...

In this paper, tube-model predictive control based on the sum of squares technique is developed for hypersonic vehicles with state-dependent input constraints. Firstly, the longitudinal non-linear model in the presence of uncertain parameters is transformed into the polytopic linear parameter varying model with bounded disturbance by feedback linearization. Then the actual input constraints are converted to the virtual state-dependent input constraints in linear multivariable polynomial. A composite feedback control law based on tube-model predictive control is designed into a convex combination of unconstrained and constrained control. The real control law can be obtained by inversion. The sum of squares technique is used to transform the polynomial constraints into the convex matrix sum of squares condition via linear matrix inequality. Finally, simulation results verify the effectiveness of the proposed controller.

The effects of external magnetic fields on the shock-wave configuration at hypersonic plasma flow field are investigated in this paper. A series of numerical simulations over various geometry configurations, namely, a blunt body and a fixed-geometry inlet forebody, have been conducted by varying the applied magnetic field under different freestream conditions. Results show that magnetohydrodynamic shock control capabilities under three types of magnetic field are ranked from weak to strong as dipole magnet, solenoid magnet, and uniform magnet field. Under the same applied magnetic field, it is easier to deflect the shock at a relatively high altitude condition, compared with the low altitude case. The bow shock standoff distance is dependent on the distribution of counter-flow Lorentz force right after shock in the stagnation region. For the oblique shock control, the function of two components of Lorentz force is different that the counter-flow one decelerates the flow and increases the shock-wave angle, while the normal one squeezes the oblique shock and deflects the streamlines.

The extremely broad flight envelope brings remarkable uncertainties to hypersonic vehicles. This paper investigates the longitudinal control problem of hypersonic vehicles with synthetical consider...

A distributed time cooperative guidance method based on the finite-time control theory is proposed to solve the salvo attack problem of multi-hypersonic vehicles in the three-dimensional space. Firstly, the second-order system is defined with the relative distance and the total leading angle as the state variables, and the finite-time controller is designed based on the multi-agent cooperative control theory for achieving the consistency of system. Novelty, taking the influence of parameters on the cooperative guidance results into account, an adaptive weight function is designed to replace the traditional constant coefficient to optimize guidance results. Then, the motion of vehicle in three-dimensional space is divided into two planes, the proportional navigation guidance (PNG) is applied in longitudinal and the lateral acceleration can be obtained via the finite-time controller and longitudinal acceleration. It is significant to establish the lateral leading angle finite-time convergence guidance law to compensate the lateral acceleration to accelerate the convergence speed. Furthermore, the biased term of terminal impact angle control is applied in PNG to achieve the terminal attack in arbitrary angle. Finally, the feasibility and robustness are verified by numerical simulations.

This study focuses on the drag reduction phenomenon caused by counterflow. Experimental studies were performed using a hypersonic shock tunnel to observe the effect of drag reduction caused by unstable counterflow. In the long penetration mode, a considerably asymmetric shock wave structure was observed. Horizontal and vertical oscillations occurred, resulting in asymmetric changes in the recirculation region when the injection pressure was low. Based on the experimental results, it was confirmed that the surface pressure decreased near recirculation regions, contributing to the drag reduction. In contrast, the occurrence of reattachment shock could result in a large increase in surface pressure. These results suggest that the drag reduction effect depends on not only the penetration length but also the flow structure. At a specific injection pressure, the Long Penetration Mode (LPM) and Short Penetration Mode (SPM) alternated; however, when the injection pressure was sufficiently high, only the SPM occurred, and a transition to a completely stable flow occurred. A maximum drag reduction of approximately 53% was observed in this hypersonic flow experiment.

The combined effect of leading-edge bluntness and backward-facing step (BFS) is critical because they affect the surface thermal loads and boundary layer transition of hypersonic vehicles, which involves engine design, passive thermal protection and aero-optical effect correction of infrared imaging window. This paper mainly studies heat flux distribution characteristics and flow structure over the backward-facing step on a flat plate with various leading-edge bluntness. The experiments were conducted in the Mach 6 hypersonic wind tunnel at various unit Reynolds numbers (from 6 × 106 to 1.8 × 107 m–1). The wall temperature was measured by temperature-sensitive paints (TSP) technique and the flow field was visualized by the nano-particles planar laser scattering (NPLS) system and schlieren technique. The results of heat flux distribution and flow structure showed that the leading-edge bluntness had little influence on recirculation region size but could effectively delay transition of the redeveloping boundary layer and reduce the surface heat flux intensity downstream of the step, and the reduction was most significant when the leading-edge diameter was equal to the step height.

Although receptivity plays a key role in the transition of hypersonic flows, most prior computational receptivity studies have neglected to study broadband frequency disturbance spectra. This work ...

The results of a parametric, numerical study of aftbody exhaust flows on a conical, airbreathing, hypersonic vehicle are presented. Three-dimensional Euler simulations were performed at Mach 9 and ...

Significantly inhomogeneous gas flow in the neighborhood of the lower surface of a plain delta wing at an angle of attack in hypersonic stream is considered. The most general case of narrow wing with the bow shock attached to the leading edge or detached from it is studied in the analytical solution of the problem obtained with the use of the asymptotic thin shock layer method. In the solution with the attached shock wave, “singular cross-sections” which are normal to the wing and have breaks of the stream surfaces due to the action of concentrated forces are found. Their input to the aerodynamic characteristics is evaluated. In flow with the detached shock, the distinctive feature of the solution relates to the formation of a thin entropy layer on the wing surface. In this layer the flow is formed under the action of the side pressure gradient. Examples of universal solutions valid for the flow description through the “external” main shock layer part and the “internal” entropy layer are constructed.

In hypersonic flows, it is often necessary to be able to trip the transition to turbulence, upstream of air intakes, for example. Direct numerical simulations have been performed to identify the ro...

Bow-shock standoff distances over sphere and spherically-blunted cone geometries were examined through experiments in two facilities capable of high-stagnation enthalpy hypersonic flows simulating Mars planetary entry conditions. High-speed and high-resolution schlieren images were obtained in the California Institute of Technology T5 reflected shock tunnel and the Hypervelocity Expansion Tube to examine facility independence of the measurements. Accompanying reacting Navier–Stokes simulations were carried out. A recently developed unified model for sphere and sphere–cone behavior was first verified for high-stagnation enthalpy CO₂ flows through simulations with thermal and chemical nonequilibrium. Shock standoff distance measurements in both facilities were found to be in good agreement with model predictions. The need to account for the divergence of the streamlines in conical nozzles was highlighted and an existing model extended to account for changes in shock curvature between parallel and conical flows. The contributions of vibrational and chemical nonequilibrium to the stagnation-line density profile were quantified using the simulation results comparing three chemical kinetic models.

The objective of this work was to investigate a multifidelity modeling approach to accurately and efficiently predict the turbulent convective heating on hypersonic inflatable aerodynamic decelerat...

With an interest in developing and studying the stability of laminar undisturbed basic-state solutions, this work is focused on accurately modeling the laminar flowfield of the boundary layer transition (BOLT) geometry under nominal and off-nominal conditions (i.e., nonzero angles of pitch and yaw). The BOLT flowfield is studied using the DPLR flow solver with MUSCL Steger–Warming fluxes using a set of five grids at different resolutions and identical grid topologies. A total of three different sets of conditions are studied: two flight conditions and one wind-tunnel-scale (33%) condition. (1) For the two sets of nominal flight operating conditions, it is found that the flow structures in the centerline region of BOLT are similar to those found in prior studies including in shape, location, and extent both vertically and spanwise, but a detailed comparison of velocity contours shows that further quantitative convergence studies are warranted. The centerline region, however, extends to at most 4 cm in semi-span at the aft end of the geometry (20% of the semi-span). Away from the centerline and where wind-tunnel-scale results have observed regions of possibly transitional behavior, the laminar flowfield converges with high accuracy. (2) For nominal wind-tunnel operating conditions, all grid resolutions simulated show good agreement in most regions as compared with prior results, with any differences falling within the scatter of existing experimental and DNS results. Aside from this focus, boundary-layer stability is examined outboard of the centerline region at nonzero pitch and yaw for a flight case, and second mode and stationary crossflow instabilities are considered. Second-mode instability is found to be locally significant at certain pitch and yaw angles particularly downstream of the swept leading edges. In addition, stationary crossflow is found to become highly amplified in significant wedges extending to the aft end of the BOLT geometry, with N-factors consistent with those found for HIFiRE-5b associated with transitional flow. The reasons for amplification of these different instabilities are also investigated from a physics-based perspective.

An experimental study on the traveling crossflow instability over a 60 $$^\circ $$ swept flat plate was conducted. The Mach number is 6, the angle of attack of the model is 5 $$^\circ $$ . The traveling crossflow waves and the secondary instability of the traveling crossflow waves were visualized by nano-tracer-based planar laser scattering (NPLS) technique. In the spanwise NPLS images, the traveling crossflow waves appeared as regular strikes, and the secondary instability appeared as small eddies attached to strikes. The wavelet transform was used to study the wavelength contents of the traveling crossflow waves. The most amplified wavelength is stable before the secondary instability happening, which is around 12 mm at $$Re_\infty $$ = 3.45 $$\times $$ 10 $$^6$$ m $$^{-1}$$ . Besides, the Reynolds number effects on the boundary layer transition and traveling crossflow instability were discussed.

The selection of an appropriate turbulence/transition model is critical when simulating the hypersonic flows based on the Reynolds-averaged Navier–Stokes (RANS) equation. In particular, a deep understanding of the validity, reliability, and limitations of existing models is an essential prerequisite to facilitate their further development. This paper reports on the assessment of two models dedicated to hypersonic boundary layer transition analysis. Both models are compared with two other models that are widely used in this field. The double ramp and shock wave laboratory scramjet intake cases are used for the validation and evaluation of the predictive performance of both models via comparison against experimental data. The results reveal that the appropriate selection of the transition model is critical to the attainment of accurate results. Moreover, the forebody of the scramjet combustion propulsion-01 (SCP-01) flight vehicle, which is a research model, is analyzed as a case study. The air intake performance of the SCP-01, as predicted by four models, is compared and analyzed at different flight altitude and Mach number conditions. The results reveal that the accuracy of the prediction of boundary layer and separated flow transitions significantly affects the flow field and corresponding air intake performance. Furthermore, the uncertainty of results obtained using the two models increases significantly with an increase in altitude. Finally, the reliability and limitations of the transition models considered in this study are examined.

Thermal protection of hypersonic vehicles is of great concern during ascent and re-entry conditions due to higher aerothermal loads. The passive forward-facing cavity is exhibited a substantial reduction of aero heating effects. The experimental studies on unsteady aspects of the flow considering the cavity geometries and heat transfer effects within the cavity are not abundant. In the present study, we investigated the heat transfer variations and nature of shock around a blunt body with cylindrical and parabolic cavity geometries. Experiments are conducted in hypersonic shock tunnels using air as the test gas for flow enthalpy conditions of 3.2 and 5.4 MJ/kg for a free-stream Mach number of 10.9 and 10.2 respectively. The cavity flow is established by verifying the measured heat flux, pressure, and shock oscillation results. The unsteady heat flux measurements and high-speed schlieren imaging is performed to study the effects. The results reveal that the flow with the parabolic cavity geometry shows a significant increase in the mean shock standoff position of about 14% and 30% compared to the blunt-nosed body, and about 7%–18% compared to the cylindrical cavity geometry, with a reduction of bow shock oscillations. As a result, a substantial reduction of heat flux about 34%–59% on the nose surface is noticed and found it is effective for mitigating the aerodynamic heating at higher flow enthalpy conditions. The flow with a cylindrical cavity geometry indicated the substantial increase in the heat flux due to the shock interaction and the significant heat flux variation in the cavity region is recognized and observed that such effects were not prevalent with the parabolic cavity geometry.

Arc-heated facilties play an important role in ground tests, such as appraisal of thermal protection systems(TPS) of various hypersonic vehicles and simulation of a high Mach scramjet's thermal structure and combustion chamber performance. This paper analyzes the requirements for developing a high power arc heater. Key technologies for developing it include high power, high voltage, high enthalpy and electrode erosion in high currents. Finally, the paper proposes the related ideas and methods for these key technologies, providing theoretical references for China's development of a super-high power arc heater.

The drag reduction and thermal protection mechanism induced by the spike and its combinations with jet strategy is investigated in the current study. A fluid-thermal coupled analysis is carried out to explore the impacts of the jet pressure ratio (PR), the length-to-diameter ratio (L/D) of the spike and the lateral jet location (DL) on the peak pressure and peak heat flux of the blunt head. The obtained results suggest that the combined spike and multi-jet strategy is the most promising technology for the drag and heat reduction. The increases in jet pressure ratios including opposing jet ratio and lateral jet ratio are advantageous and provide the maximum reduction in the drag and heat flux. Specifically, for the cases of PRo = 0.2 and PRo = 0.6, the peak heat fluxes are decreased by about 2.25 kW/m2 and 0.37 kW/m2 respectively after 0.6 s’, and the decrease rate becomes more slowly during the coupling process. Further, the negative thrust produced by the opposing jet plays a key role in the total drag component. For the L/D effect, the peak pressure and drag coefficient decreases when L/D increases. However, it is noteworthy that when L/D = 0.5, the peak heat flux is close to that when L/D = 2, which, is at a relatively low level of 104.83 kW/m2. In addition, the lateral jet location has a decisive influence on the flow structure, drag coefficient and heat flux distribution, and DL = 8 mm gives the minimum drag and heat flux while the peak pressure is the maximum.

The study considers the solution of a number of problems of supersonic and hypersonic gas dynamics using a model that takes into account the dissociation and ionization of air. The results of verification and validation of the developed numerical method using various difference schemes (the Roe scheme, Rusanov scheme, AUSM scheme) for discretizing convective flows are presented. The formulation of the mathematical model for high-temperature air uses the presence of equilibrium chemical reactions of dissociation and ionization. For this purpose, at high incoming flow velocities, the Kraiko model is applied, which includes equilibrium chemical reactions in air at high temperatures. To discretize the basic equations, the finite volume method on an unstructured grid is applied. One of the features of the constructed mathematical model is the implementation of the transition between physical and conservative variables. Relationships are given, with the help of which the transition from conservative variables to physical ones and vice versa is carried out when using the high-temperature air model. To ensure the stability of numerical calculations, an entropy correction is introduced. The decrease in entropy in the solution of hyperbolic equations is excluded by introducing an artificial viscosity according to Neumann, as well as by using the Godunov method with an exact solution of the Riemann problem and methods based on the approximate solution of the problem of the decay of an arbitrary discontinuity. A number of problems of supersonic gas dynamics (supersonic flow in a channel with a straight step and supersonic flow around a sphere) are numerically solved taking into account high-temperature effects. The criteria for the accuracy of numerical calculations related to the location of shock-wave structures are discussed. The calculated shock-wave structure of the flow is compared with the data available in the literature, as well as with calculations using the perfect gas model. Some results of numerical calculations are compared with the available experimental data. The shock-wave flow patterns obtained in the framework of the inviscid model, which takes into account the effect of viscosity and its dependence on temperature, and the turbulent flow model are compared. On the basis of numerical simulation data, the influence of viscous effects on the flow characteristics in a channel with a straight step and hypersonic flow around a sphere is considered. The influence of various numerical factors on the shape of the bow shock and the presence of fluctuations in the solution behind the shock is emphasized. As part of the work, a computational module was prepared for the commercial package Ansys Fluent, implemented with the help of user programming tools. The prepared module expands the standard capabilities of commercial software focused on solving computational gas dynamics problems, and is available to Ansys Fluent users for solving hypersonic aerodynamics problems. The developed means of numerical simulation can be useful in the design and optimization of hypersonic aircraft.

A novel inhomogeneous planar layer variable thickness radome (VTR) is presented in this letter. To improve the performance of the radome, the effect of thermal coefficient of expansion on radome wall thickness is incorporated in the design stage. The proposed radome exhibits superior power transmission compared to the conventional monolithic constant thickness radome and the monolithic VTR. A Geometrical optics based 3-D ray tracing with aperture integration technique is used for the performance prediction of the radome. The designed radome exhibits an insertion loss as low as 0.006 dB and a maximum boresight error of 2.4 mrad. A time lapse study on the radome performance due to varying temperature distribution as the radome mounted vehicle attains hypersonic velocity is also presented. Superior performance at hypersonic environment makes this radome an attractive choice for airborne applications.

The nonlinear analyses of the hypersonic and high-enthalpy boundary-layer transition had received little attention compared with the widely-studied linear instabilities. In this work, the oblique-mode breakdown, as one of the most available transition mechanisms, is studied using the nonlinear parabolized stability equations (NPSE) with consideration of the thermal-chemical non-equilibrium effects. The flow over a blunt cone is computed at a free-stream Mach-number of 15. The rope-like structures and the spontaneous radiation of sound waves are observed in the schlieren-like picture. It is also illustrated that the disturbances of the species mass and vibrational temperature near the wall are mainly generated by the product term of the wall-normal velocity disturbance and the mean-flow gradient. In comparison to the CPG flow, the TCNE effects destabilize the second mode and push upstream the N factor envelope. The higher growth rate of the oblique wave leads to stronger growth of the streamwise vortices and harmonic waves.

A comprehensive review of all relevant experimental data was completed, including recent data for the drag coefficient for a sphere in supersonic and hypersonic flows. The primary characterization ...

This two-parts article introduces a novel hybrid propulsion system based on the DEA compressor. The system encompasses a Pulse Detonation TurboDEA as the master engine that supplies several full-electric ancillary thrusters called DEAThruster. The system, called the propulsion set, can be categorized as a distributed propulsion system based on the design mission and number of ancillary thrusters. Part B of this article explains the performance sizing of the propulsion set designed in part A. Evaluating the performance of the propulsion, computer programs are written for all major components of the both master engine and ancillary thruster. The intake, compressor, detonation process, diffusers, axial turbine, and exit nozzle are modeled under certain flight conditions, and their performances are revealed and analyzed. The flight conditions are considered from the static condition at the sea level up to flight Mach number 5 at an altitude of 20,000 m. The performance of the propulsion set is also compared with some aircraft propulsions modeled by similar studies in all important aspects.

In this paper, the restart processes of two generic rectangular hypersonic inlets with different internal contraction ratios have been experimentally studied under a freestream Mach number of 5.0. ...

Since proposed, the self-similarity variables based genuinely multidimensional Riemann solver is attracting more attentions due to its high resolution in multidimensional complex flows. However, it needs numerous logical operations in supersonic cases, which limit the method’s applicability in engineering problems greatly. In order to overcome this defect, a hybrid multidimensional Riemann solver, called HMTHS (Hybrid of MulTv and multidimensional HLL scheme based on Self-similar structures), is proposed. It simulates the strongly interacting zone by adopting the MHLLES (Multidimensional Harten-Lax-van Leer-Eifeldt scheme based on Self-similar structures) scheme at subsonic speeds, which is with a high resolution by considering the second moment in the similarity variables. Also, it adopts the MULTV (Multidimensional Toro and Vasquez) scheme, which is with a high resolution in capturing discontinuities, to simulate the flux at supersonic speeds. Systematic numerical experiments, including both one-dimensional cases and two-dimensional cases, are conducted. One-dimensional moving contact discontinuity case and sod shock tube case suggest that HMTHS can accurately capture one-dimensional expansion waves, shock waves, and linear contact discontinuities. Two-dimensional cases, such as the double Mach reflection case, the supersonic shock / boundary layer interaction case, the hypersonic flow over the cylinder case, and the hypersonic viscous flow over the double-ellipsoid case, indicate that the HMTHS scheme is with a high resolution in simulating multidimensional complex flows. Therefore, it is promising to be widely applied in both scholar and engineering areas.

Supersonic accelerators use sabots to suppress bore balloting. However, sabot separation induces flight deviation in projectiles owing to their interactions with fluids and shock waves. A seamless ...

This work focuses on the heat flux characteristics of a hypersonic blunt body with a tangential Mach 3 supersonic cooling film. Experiments were carried out with main flow’s Mach numbers of 6, 7, a...

In this work, the stability of Mach 20 flows past a 6 deg wedge in thermal–chemical nonequilibrium (TCNE) is studied by means of linear parabolized stability equations (PSEs) in combination with ad...

Experiments were conducted on a cylinder-induced shock-wave/boundary-layer interaction (SBLI) at Mach 6. The state of the interaction was brought through transition using a combination of boundary-...

Three of the key associative ionization reactions that lead to plasma formation around hypersonic vehicles are studied in detail. For production of NO+, O2+, and CO+, cross sections derived from ex...

Flutter of unmanned aircraft at hypersonic speeds is one of the understudied problems facing the designers of hypersonic vehicles. Modern methods for calculating aeroelastic stability either solve simplified versions of real problems (for example, without taking physicochemical phenomena into account) or require high computing power. This paper describes a methodology for calculating the supersonic and hypersonic flutter of an aircraft using standard engineering software and additionally developed software modules. The justification of the need to refine the existing methods for calculating the aircraft aeroelasticity taking into account the real geometry of the structure and with the possibility of accounting for physicochemical processes occurring in the air during the movement of bodies at high speed is provided. The theoretical principles of calculating aeroelastic stability taking into account these factors are developed, and three examples of calculating the aeroelastic stability of model objects are given.

The objective of the present paper is to reveal the intrinsic mechanisms by which two-dimensional surface indentations influence the laminar–turbulent transition of natural route in hypersonic boun...

Waveriders have a high lift-to-drag ratio in hypersonic states, while the flexibility of design methods must still be improved because of some deficiencies such as unsatisfactory off-design perform...

This article presents a fixed-time accommodation strategy of actuator faults for hypersonic gliding vehicles (HGVs). The approach against actuator faults is incorporated by sliding mode control (SMC), bilimit homogeneity, and adaptive techniques, with consideration of practical constraints and model uncertainties. The resulting fault-tolerant attitude control law allows the fault compensation to be completed in a fixed time, in view of the limited time available for recovery of a faulty HGV. The effectiveness of the presented scheme is validated by comparing it to the finite-time fault accommodation scheme. Note to Practitioners —Hypersonic vehicles have drawn significant interests due to the features of super-fast and flexible maneuverability. The use of advanced fault-tolerant control (FTC) techniques is expected to guarantee safety during hypersonic vehicle operation. A challenging problem is how to counteract actuator faults promptly and effectively in the presence of practical constraints and model uncertainties. This article presents a fixed-time FTC algorithm for a hypersonic gliding vehicle.

The rapid development of hypersonic vehicles has motivated the related research dramatically while the evasion of the hypersonic vehicles becomes one of the challenging issues. Different from the work based on the premise that the pursuers’ information is fully known, in this paper the evasion guidance for air-breathing hypersonic vehicles (AHVs) against unknown pursuer dynamics is studied. The gradient descent is employed for parameter estimation of the unknown dynamics of the pursuer. The energy-optimized evasion guidance algorithm is further developed by taking the acceleration constraint and energy optimization into consideration. Under the proposed algorithm, the system can deal with the unknown pursuer dynamics effectively and provide more practical guidance for the evasion process. The simulation results show that the proposed method can enable the AHV to achieve successful evasion.

Lateral–directional stability is a critical issue for the design of any kind of aircraft. For the hypersonic waverider, the nonaxisymmetric, flat, and slender geometric feature makes it susceptible...

Based on non-affine models of hypersonic space vehicles, the tracking control problem of hypersonic vehicles is studied and analyzed in this article using funnel robust model-free control mechanism considering parametric uncertainty and external disturbances. First, the control system is decomposed into altitude subsystem and velocity subsystem. For altitude subsystem, we propose a concise funnel robust model-free control mechanism based on error driving, and a novel model transformation approach is applied to the controller design. The new model-free controller only contains a Hurwitz stable term and a filtering term, and does not need precise motion model and too much calculation, so it can improve the calculation speed of the system. For velocity subsystem, only a concise proportional-integral controller is needed to meet the tracking requirements. Moreover, the devised controller is capable of guaranteeing funnel performance on the altitude and velocity tracking errors. Finally, numerical simulation results are presented to verify the efficiency of the design.

The plasma sheath, which is the weakly ionized gas layer that envelopes hypersonic vehicles in near space, could shield communication signals and yield communication blackout. In the recent couple of decade, the extremely high frequency (EHF) communication is being considered as a potential solution to the blackout problem. On the other hand, previous studies used to focus on the EHF wave propagation in idealized plasma slabs or plasma sheaths covering blunt-coned vehicles. In the present letter, the plasma sheath enveloping sharp-coned hypersonic vehicles was modeled. The characteristics of EHF wave propagation in the sharp-coned plasma sheath were investigated. According to the study, the maximum electron densities and electron collision frequencies in sharp-coned plasma sheath is in the same magnitude of that in blunt-coned plasma sheath. However, the thickness of sharp-coned plasma sheath is much lower than that of blunt-coned plasma sheath. In addition, the power transmission rate at the frequencies of 140 and 225 GHz is high enough to afford communication tasks. Moreover, the reflection plays important roles in the energy loss of EHF wave propagation in sharp-coned plasma sheath, whereas the reflection in blunt-coned plasma sheath is ignorable. According to the analysis, the reason for such difference is that the gradient of relative permittivity in sharp-coned plasma sheath is larger than that of blunt-coned plasma sheath, which yields significant reflection to the EHF waves.