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This paper focuses on the design of fuzzy secure control for a class of coupled systems, which are modeled by a nonlinear $N$ -dimensional (N-D) parabolic partial differential equation (PDE) subsystem and an ordinary differential equation (ODE) subsystem. Under stochastic deception attacks, a fuzzy secure control scheme is designed, which is effective to tolerate the attacks and ensure the desired performance for the considered systems. A new fuzzy-dependent Poincare-Wirtinger's inequality (PWI) is proposed. Compared with the traditional Poincare's inequality, the fuzzy-dependent PWI is more flexible and less conservative. Meanwhile, an augmented Lyapunov-Krasovskii functional (LKF) is newly constructed, which strengthens the correlations of the PDE subsystem and ODE subsystem. Then, on the ground of the fuzzy-dependent PWI and the augmented LKF, new exponential stabilization criteria are set up for the PDE-ODE coupled systems. Finally, a hypersonic rocket car is presented to verify the effectiveness and less conservatism of the obtained results.

In this paper, the tracking control problem of non-minimum phase flexible air-breathing hypersonic vehicles (AHSV) is investigated subject to actuator fault, external disturbances and parameters uncertainties. The study is began with a series of control-oriented manipulations: first, the input–output dynamics are derived by using feedback linearization method and the internal dynamics of AHSV are constructed; then, the zero dynamics stability analysis is conducted to verify the non-minimum phase characteristic of AHSV. In order to realize output tracking of the non-minimum phase system with sufficient accuracy, an adaptive fault tolerant controller (FTC) is proposed based on an output-redefinition making the zero-dynamics with respect to the new output stable. Additionally, robust adaptive laws are utilized for the estimation of unknown parameters and actuator failure compensation of the AHSV model. Furthermore, the stability of the closed-loop system is analyzed based on the Lyapunov stability theory. At last, the numerical simulation results are provided to demonstrate the effective tracking performance of the proposed FTC scheme.

Air-breathing hypersonic vehicle has great military and potential economic value due to its characteristics: high velocity, long range, quick response. Therefore, the development of hypersonic vehicle and its guidance and control technology are reviewed in this paper. Firstly, the development and classification of hypersonic vehicles around the world are summarized, and the geometric configuration and mission profile of typical air-breathing hypersonic vehicle are given. Secondly, the control difficulties of air-breathing hypersonic vehicle are introduced, including integrated design of engine and fuselage, static instability, strong nonlinearity, uncertain aerodynamic parameters, etc. According to its control requirements, the control methods considering external disturbance, fault-tolerant control methods, anti-saturation methods, and prescribed performance control methods considering transient performance constraints are summarized respectively. The classification and comparison of various control methods are given, and the frontiers of theoretical development are analyzed. Finally, considering the effects of composite disturbances, the design of terminal guidance law under multiple constraints is overviewed, including guidance law with angle constraint, velocity constraint, acceleration constraint and time constraint. Similarly, the classification of guidance law design methods under different constraints, their advantages as well as the future development trend and requirements are introduced.

The present study completes the simulation of a Terminal High Altitude Area Defense–like missile’s lateral jet flowfield at a rarefied high-altitude atmosphere using the in-house parallel direct si...

A sliding mode based fault-tolerant control method using neural learning is studied for hypersonic reentry vehicle (HRV) in this paper. Based on the non-singular second-order terminal sliding mode, the composite neural learning is adopted to deal with the system uncertainties caused by the additive faults. The main point is to construct the prediction error to evaluate the performance of intelligent approximation. Meanwhile, a disturbance observer is employed to estimate the unknown disturbance while the actuator multiplicative fault is compensated with an adaptive law. The simulation of HRV is conducted to verify the effectiveness of the proposed fault-tolerant control scheme.

The design of optimal guidance law for intercepting a near-space hypersonic maneuvering target with bounded inputs is considered. Firstly, a maneuvering model for near-space hypersonic aircraft is given. Then, the aircraft acceleration prediction can be obtained using this model with two neural networks. By using the target acceleration prediction, which is taken into account when calculating the Zero Effort Miss (ZEM), an optimal sliding-mode guidance law is proposed to fulfill the guidance task. An adaptive sliding-mode switch term is designed to deal with actuator saturation and prediction errors. Finally, numerical simulations show that the proposed guidance law can reduce the energy consumption and the terminal acceleration command of the interceptor effectively.

Estimating heat-flux peak location and magnitude is important to efficiently design thermal protection systems of hypersonic vehicles, which are characterized by strict low-weight requirements. Thi...

Pressure-gain combustion has gained attention for airbreathing ramjet engine applications owing to its better thermodynamic efficiency and fuel consumption rate. In contrast with traditional detonation induced by a single wedge, the present study considers oblique shock interactions attached to double wedges in a hypersonic combustible flow. The temperature/pressure increases sharply across the interaction zone that initiates an exothermic reaction, finally resulting in an Oblique Detonation Wave (ODW). Compared with the case for a single-wedge ODW, the double-wedge geometry has great potential to control the initiation of the ODW. As a tentative study, two-dimensional compressible Euler equations with a two-step induction-reaction kinetic model are used to solve the detonation dynamics triggered by a double wedge. The effects of the wedge angles and wedge corner locations on the initiation structures are investigated numerically. The results show an ODW complex comprising three Oblique Shock Waves (OSWs), an induction zone, a curved detonation front, and an unburned/low-temperature gas belt close to the surface of the second wedge. Both the increasing wedge angle and downstream wedge corner location lead to an abrupt OSW-ODW transition type, whereas the former corresponds to the shock–shock interaction and the later has a greater effect on the exothermic chemical process. Analysis of the shock polar and flow scale confirms that the OSW-ODW initiation structure mainly depends on the coupling of shocks and heat release in a confined initiation zone.

Analysis of multidimensional in-depth thermal response for the phenolic impregnated carbon ablator (PICA)/CV-1144-0 material system in NASA Langley Research Center Hypersonic Materials Environmenta...

This paper develops a decentralized actuator fault compensation scheme for hypersonic vehicles, using a model transform based backstepping control design. Different from the actuator dynamics represented by second-order ordinary differential equations (ODE) studied in the existing literature, a new actuator dynamics is studied in this paper, modeled by a fourth-order partial differential equation (PDE), i.e., Euler-Bernoulli beam (EBB) equation. An invertible model transformation is introduced to transform the EBB PDE-ODE cascade system into a heat-like PDE-ODE cascade system, and a new target system is constructed from a two-step backstepping procedure. A novel decentralized fault-tolerant control scheme is proposed for the heat-like PDE-ODE cascade system and the well-posedness of the system solutions is analyzed. The closed-loop system stability and the well-posedness property of the original EBB PDE-ODE cascade system are ensured. Simulation results illustrate the effectiveness of the proposed fault compensation scheme.

The effect of thermochemical kinetics modeling on hypersonic flow over a double-cone geometry is investigated. The double-cone is simulated using three different approaches based on the Park model:...

Experimental investigations of two interacting spheres were carried out in a Mach 4 flow with a static pressure of 2.666 Pa. This study was performed in the MARHy wind tunnel, an adjustable supersonic or hypersonic low-density facility. The main purpose is to estimate how the interactions between spherical debris modify the aerodynamic forces during their atmospheric entry. Indeed, around 80 km, altitude experimentally reproduced thanks to MARHy flow conditions, most debris fragment. The multiple debris, resulting from fragmentation, interact with each other’s, which may impact their trajectory. The purpose of this work is to analyze the behavior of a moving sphere placed in the wake of another one. The moving sphere is suspended with two thin wires, so it can almost freely move in the flow direction and physically respond to the acting flow forces of the wake of the first one. Aerodynamic forces are then calculated from the sphere angle deflection. The qualitative analysis of the experimental results will be discriminated first with respect to the trajectory of the second sphere, and secondly with respect to the evolution of the aerodynamic forces according to the sphere.

Plasma generated by hypersonic aerocraft is one of the research subjects in computational electromagnetics. When the aerocraft flies at an ultra-high speed in the near space, plasma sheath will be generated in the surroundings and it will trigger varying changes in aerocraft radar electromagnetic scattering properties. Plasma sheath brings about tremendous impacts on the communication between the aerocraft and the outside world, and even ends with "blackout effects" to interrupt communication in severe cases. The interplay between the plasma sheath and the electromagnetic wave involves multiple disciplines covering aerodynamics, plasma sheath physics, and electromagnetic field theory. Therefore, it is challengeable to build a rigorous numerical modeling of the plasma sheath. In this paper, the latest technologies in the electromagnetic simulation of hypersonic aerocraft with plasma sheath are reviewed. Applied cases review and illustrate extensive approaches from approximate calculations to accurate calculations, including analytical methods, differential equation methods, integral equation methods and high-frequency approximation methods. The paper assesses the RCS of electromagnetic wave in analyzing the electromagnetic scattering problem in plasma. With the considerations the reality of electromagnetic simulation and future challenges, the paper talks about the efficiency, advantages and disadvantages of these approaches in simulating hypersonic aerocraft targets.

Shock tube experiments are a primary means of obtaining ground test data for the hypersonic regime. Accurate characterization of the test gas is crucial to understanding experimental results. However, characterization of the flows produced behind the shock wave has historically proven challenging. This paper applies a methodology to calculate the shocked test gas properties using the experimentally recovered shock speed profile. Static pressure, pitot pressure, and heat transfer predictions are found to closely match the experimental data for a range of shock trajectories with both argon and air test gases. Thermochemical variations in the test gases are found to depend strongly on variations in shock speed along the tube. It is shown that characterization of the test gases requires accommodating the influence of wave effects associated with the varying shock speed. Tube diameter is found to influence test time significantly in interaction with a variable shock speed, and also the magnitude of nonuniformities in the test gas. Location and number of shock timing stations in experimental facilities are found to play a vital role in the ability to accurately characterize the test gas of a given experiment.

The discontinuity at the thermal protection system tiles covered on an aerospace vehicle is modeled as an open cavity to numerically investigate rarefied hypersonic flows over it under various flow control styles using the direct simulation Monte Carlo method. The influence of flow control on flow structures (e.g., shock wave and vortex) and aerothermal properties (e.g., temperature distribution and wall heat flux) is analyzed in detail. The flow structures and aerothermal properties of the cavity flowfield are both changed greatly by the square control block on the cavity opening line, such as a reduction of about 79.5% of the maximum of wall heat fluxes near the reattachment corner, while these changes are gradually weakened with the control block moving down. The control blocks with different shapes still have the ability to change flow structures and thus the heat flux coefficients over the cavity surfaces, but shape has little effect, and the square control block has the strongest wall heat flux compared with the circular, triangular, and wedged control blocks. For the supersonic jet with different directions, the leftward jet has the best performance in eliminating high temperature gas inside the cavity and over the downstream wall and reducing wall heat fluxes over cavity surfaces, followed by the downward jet. The main reason is that the leftward and downward jets can effectively prevent momentum and energy exchanges between freestream with cavity surfaces.

Hypersonic high-enthalpy wind tunnels have been a challenge to ground tests in aerospace research area for decades and its test flow uniformity is one of the most important parameters for evaluating wind tunnel performances. Regarding to the performance requirement, theories and methods for designing hypersonic flow nozzles at high enthalpy conditions are quite difficult, but very interesting topics, especially when air molecule dissociations take place in wind tunnel test gas reservoirs. In this paper, fundamental theories and important methods for nozzle designs are briefly reviewed with the emphasis on two-dimensional axisymmetric nozzles for hypersonic high-enthalpy wind tunnels, including the Method of Characteristics (MOC), the graphic design method, the Sivells method, the theory for boundary correction, and the CFD-based design optimization methods. These theories and methods had been proposed based on several physical issues, respectively, which play important roles in nozzle flow expansion processes. These issues cover the expansion wave generation and reflection, the boundary layer development, the real gas effect of hypersonic high-enthalpy flows. Difficulties arising from applications of these methods in hypersonic high-enthalpy nozzle design are discussed in detail and the state of the art of the nozzle design technologies that have reached for decades is summarized with some brief comments. Finally, the prospect for the hypersonic nozzle design methods, and its numerical and experimental verifications are provided with from authors’ viewpoint for readers’ reference.

This review paper examines ten current ceramic radome materials under research and development and provides a comprehensive overview of available high temperature and high frequency data from literature. An examination of metamaterials for radio-frequency transparent radomes is given and our preliminary experimental results of a high-temperature metamaterial design are presented. The next-generation hypersonic vehicles’ radome temperatures will exceed 1000℃ and speeds will exceed Mach 5. An ideal radome material will have a high flexural strength, low dielectric constant and loss tangent, and high resistance to thermal shock and corrosion. The microstructural effect on the dielectric and mechanical properties and the effects of environmental factors such as rain are discussed. The impact of metamaterial structure on key radome factors such as boresight error, gain, and polarization is examined. After examining the associated benefits with the use of metamaterials, our preliminary results for a potential high-temperature metamaterial design are presented.

With the increasing altitude and speed of the hypersonic vehicles, the temperature of the combustion chamber wall must be substantially decreased to ensure thermal and structural reliabilities. As a great supplement to the main regenerative cooling method without additional coolant, film cooling with onboard hydrocarbon fuel is widely used to decrease the wall heat flux. In this paper, a typical cylindrical film injection hole is employed to study the flow and heat transfer characteristics of gaseous film cooling with hydrocarbon fuel in the supersonic main flow, and the effects of transverse-non-uniform distributions of blowing ratio and injection angle are analyzed in detail. The results show that entrainment process between cooling films with different blowing ratios and injection angles would influence the film cooing efficiency to different degree, and proper combination of these two factors can improve the film cooling efficiency and decrease the flow rate of the cooling film. And the empirical correlation of supersonic hydrocarbon film cooling efficiency is established for the uniform blowing ratio cases. For the transverse non-uniform blowing ratio, the overall cooling efficiency decreases due to the spanwise and height-wise ejections between the two groups of cooling film. And for the cases with combination of transversely non-uniform blowing ratio and unequal injection angle, the film cooling effectiveness per blowing ratio would increase by 8.0%.

This letter proposes a modified polar format algorithm (PFA) applied to the highly squinted synthetic aperture radar (SAR) onboard an accelerating hypersonic platform. In conventional PFA, the range wavenumber is orthogonally decomposed to obtain the 2-D wavenumbers corresponding to the imaging Cartesian coordinates. This operation, however, is inadequate when facing with more complex motion trajectory. In this letter, the essence of the proposed method is to homogenize the spatially nonuniformly sampled echoes by combining interpolation with the generalized wavenumber definition. The imaging polar coordinate system in the slant plane provides new space-wavenumber Fourier transform pair, and the 1-D interpolation compensates for the range-variant phase errors. The proposed algorithm, thus, enables fast and effective imaging of hypersonic SAR in large squint angles and high maneuverability dive mode. Its performance is discussed and analyzed in this letter, including the resolution and the size of the imaging scene. The effectiveness, superiority, and application value of the proposed algorithm are verified by simulation.

Among flourishing control studies on air-breathing hypersonic vehicles (AHVs), actuator characteristics have not been fully accommodated. This study gives a comprehensive consideration to physical ...

A numerical simulation study of laminar hypersonic cold airflow over the backward-facing step (BFS) is performed at a freestream Knudsen number of 0.00009589 for resolving the flow field. The effects on the wall properties caused by the expansion of the flow and the separation of the boundary layer at the step corner have been investigated. The contributions of pressure and viscosity to the coefficients of lift and drag have been elucidated. It is seen that an increase in step height “nh” from n = 1 to 5 for a freestream flow at Mach 7.6 produces an incremental effect on the CL/CD ratio, whose range is between −1 to 1 for the variations in the step height simulated. The current study concludes that BFS yet remains a geometry that contributes to more drag than lift, however that effect can be minimized by increasing the step height during the design process of hypersonic vehicles for the above-mentioned Mach number.

A novel technique was proposed in this work to generate hypersonic water vapor plasmas in an inductively coupled plasma wind tunnel. Based on it, the oxidation behavior of a typical SiC coating was studied at varying heat flux in reduced pressure. The key observation is an accelerated corrosion of the SiC coating at

A novel optical-tracking technique is presented for aerodynamic measurements on free-flying models in hypersonic wind tunnels. This technique is a generalization of the edge-fitting concept develop...

Shock wave–boundary layer interaction (SWBLI) is one of the most prevalent challenges in the field of fluid mechanics. Excessive interaction may lead to strong flow separation, which is an important factor causing the deterioration of aircraft performance. Reynolds averaged Navier–Stokes (RANS) models remain the primary methods for predicting flow separation for engineering applications. RANS simulation results typically have uncertainty owing to the drawbacks of the model structure and the variability of the parameters. Therefore, evaluation of a turbulence model based on nominal predictions may be insufficient for reflecting its actual simulation capabilities. In this study, the Bayesian method is employed to re-evaluate the simulation capabilities of three k–ω models for hypersonic SWBLI flows: the original k–ω model and two k–ω models with compressibility corrections (Wilcox's and Zeman's corrections). This is achieved by perturbing the model parameters and combining with experimental data. Bayesian model evaluation based on different datasets is performed using the embedded model error method, and the three turbulence models are compared from three aspects: posterior mean of the quantities of interests, model error, and Bayesian evidence. The results show that the Bayesian calibration can effectively improve the predictive abilities of the three models for wall pressure and wall heat flux. Different from the large differences in the nominal results, the performance of the two calibrated k–ω models with compressibility corrections are very close and superior to that of the original model. Moreover, the evaluation results of this study are also proven to be applicable to other similar flows.

The aerodynamic shape of the hypersonic forebody/inlet integrated vehicle is usually designed according to the inviscid flow. However, due to the effect of air viscosity, there is a boundary layer flow near the wall surface. In order to get the actual flow characteristics and performance parameters of the inviscid design vehicle close to the inviscid design values, the inviscid design surface must be corrected by considering the viscous boundary layer. In this paper, the theory and method of the boundary-layer viscous correction (BLVC) for the forebody/inlet integrated configuration are studied. Firstly, two modification schemes of the BLVC for the forebody/inlet integrated configuration are proposed. The calculation method of boundary-layer displacement thickness (BLDT) based on the momentum integral equation of steady two-dimensional (2D) axisymmetric compressible flow is described in detail, and two calculation methods of the BLDT are compared and analyzed. Secondly, the numerical simulation method used in this paper and its validation are introduced. Thirdly, an axisymmetric configuration of the forebody/inlet integration is selected as the first case, and two modification schemes of BLVC are studied and analyzed by numerical simulation. The effects of the two modification schemes of BLVC on the internal and external flow characteristics and aerodynamic performance of the axisymmetric configuration of the forebody/inlet integration are compared and analyzed. Finally, a full-waverider configuration of the forebody/inlet integration is selected as the second case, and the second modification scheme of the BLVC is used to modify the full-waverider configuration. The shock wave characteristics and inlet performance parameters of the modified full-waverider configuration are studied by both the numerical simulation and the wind tunnel experiment. The obtained results prove that the calculation method of the BLDT by a numerical integration method to solve the momentum integral equation of the steady 2D axisymmetric compressible flow is correct and the BLVC method and scheme for the forebody/inlet integrated configuration are effective. The modification scheme 1 of the BLVC can only improve the characteristics of forebody shock wave incident on the inlet lip under viscous conditions, that is, to improve the performance of inlet to capture precompressed air flow, while the modification scheme 2 of the BLVC can simultaneously improve the characteristics of forebody shock wave incident on the inlet lip and the characteristics of the lip shock wave cancelled on the shoulder of the inlet under viscous conditions, and make the air flow capture performance of the inlet and the pressure rise ratio at the inlet exit close to those of the inviscid design state.

Hypersonic boundary layers are crucial in aerospace applications such as hypersonic glide vehicles, rockets, and other advanced space vehicles. Hypersonic flows present unique transport phenomena, ...

The ability to measure the very high heat fluxes that typically occur during the hypersonic re-entry phase of space vehicles is generally considered a subject of great importance in the aerospace field. Most of the sensors used for these measurements need to be checked periodically and re-calibrated accordingly. Another bottleneck relates to the need to procure thermal sources that are able to generate reliable reference heat fluxes in the range between 100 and 1000 kW/m2 (as order of magnitude). In the present study, a method is presented by which, starting from a calibration system with a capacity of approximately 500 kW/m2 only, heat fluxes in the range of interest for hypersonic applications are generated. The related procedure takes advantage of established standards for the characterization of a radiative heat flux. It also builds on the hybrid radiative-convective nature of typical hypersonic heat fluxes and the yet poorly explored possibility to use convective sources of heat to produce high-intensity fluxes. The reliability of such a strategy has been tested using a high enthalpy supersonic flow facility relying on an electric arc-heater and pure Nitrogen as work gas. Stagnation-point heat fluxes have been successfully measured (with reasonable accuracy) in the range between 600 and 1500 kW/m2 for values of the centerline enthalpy spanning the interval from to 6 to 24 MJ/kg.

To track a non-cooperative hypersonic glide vehicle (HGV) without any precise information, an approach to the state estimation is presented based on a robust UKF-based filter (RUKFBF) in this paper. The HGV has an uncertain reentry motion because of unknown maneuvers which is a primary factor leading to degradation of tracking accuracy. Aiming at enhancing accuracy, the strong tracking algorithm (STA) is introduced to addressing the model error caused by a bank-reversal maneuver of HGV. Furthermore, the Huber technique is employed to deal with possible measurement model errors. In the RUKFBF, mutual interferences are suppressed between the STA and the Huber technique via two strategies. The one is that the calculation of the fading factor in the STA adopts an unmodified measurement noise covariance, and the other one is that two judgment criteria are proposed to limit large fading factors in the presence of measurement model errors. To simulate real tracking scenarios, the RUKFBF is tested through tracking a HGV trajectory considering a practical guidance strategy. Simulation results demonstrate the effectiveness of the RUKFBF in the presence of model errors and the observability of the estimated state.

The supersonic mode in hypersonic flows is studied numerically. The radiation characteristics of this mode in a Ma=6 flat-plate boundary layer with cooled walls are focused on. The linear stability...

During atmospheric reentry at hypersonic speed and during very-high-speed flight in the atmosphere, vehicles experience extremely high thermal loads for relatively long time. To handle this problem...

Scramjet engines are one of the most promising hypersonic airbreathing propulsion technologies for robust, efficient, and economical access to space. Multi-objective design optimization has been co...

SparkJet (SPJ) actuator was used to control shock and shock-on-shock interaction in high-enthalpy hypersonic flow with a total temperature of 4490 K and Mach number 6.9 for the first time. Aiming at the characteristics of rarefied gas in hypersonic flight environment, the control authority of SPJ actuator was improved by air supply method. The process of SPJ interacting with high-enthalpy hypersonic crossflow near a ramp was obtained by using high-speed schlieren camera. When pressurized cavity pressure was 133 kPa, ramp distance was 22 mm and ramp angle was 50, SPJ can eliminate 68.3 % of the whole ramp shock. Results of different experiment cases showed that, within a certain range, the control effect of SPJ on ramp shock improves with the increase of pressurized cavity pressure, ramp distance and the decrease of ramp angle. The elimination of shock by SPJ can be explained by the upward motion of the sonic line in the boundary layer due to the formation of SPJ shock, the thickening of boundary layer and the heating of local flow field. Except for the elimination of shock, the shock intersection point moves upstream and downstream corresponding to the strength changes of SPJ and SPJ shock. Accordingly, local high heat flow around the shock intersection point can be decreased. The control effect of SPJ on shock-on-shock interaction can be improved by properly increasing discharge energy.

In the design of the hypersonic airliner which can greatly shorten the flight time and conduct space travel, it is of great significance for a Riemann solver to accurately simulate hypersonic flows. In this survey, the research process on the Riemann solver for hypersonic flows is reviewed, including the constructions of the traditional Riemann solvers, improvements of the traditional Riemann solvers for the shock anomaly, and the importance of the all-speed Riemann solver for hypersonic flows. Moreover, constructions and applications of the all-speed Riemann solvers are presented. It is obvious that the Riemann solver should be capable of obtaining physical solutions at low speeds and be robust against shock anomaly near strong shocks at the same time in simulating hypersonic flows.

To understand fundamental problems in hypersonic laminar-turbulent boundary layer transition for three-dimensional complex vehicles, a new standard model with typical lifting-body features has been proposed, named as hypersonic transition research vehicle (HyTRV). The configuration of HyTRV is fully analytical, and details of the design process are discussed in this study. The transition characteristics for HyTRV are investigated using three combined methods, i.e., theoretical analyses, numerical simulations, and wind tunnel experiments. Results show that the fully analytic parameterization design of HyTRV can satisfy the model simplification requirements from both numerical simulations and wind tunnel experiments. Meanwhile, the flow field of HyTRV reveals typical transition mechanisms in six relatively separated regions, including the streamwise vortex instability, crossflow instability, secondary instability, and attachment-line instability. Therefore, the proposed HyTRV model is valuable for fundamental researches in hypersonic boundary layer transition. A new standard model with typical lifting-body features for boundary layer transition research has been proposed. Results of three combined analysis methods show that the flow field of HyTRV reveals typical transition mechanisms in six relatively separated regions, including the streamwise vortex instability, crossflow instability, secondary instability, attachment-line instability. Therefore, the proposed HyTRV model is valuable for fundamental researches in hypersonic boundary layer transition.

Based on the new 2A' and 4A' potential energy surfaces of NO2 fitted by Varga et al., we conducted a quasi-classical trajectory study on the N(4S) +O2(X3Σg- ) → NO(2Π) + O(3P) reaction, focusing on the high vibrational state up to ν = 25. For different rovibrational states of O2, within the relative translational energy (Ec) range of 0.1-30 eV, the total exchange cross section (ECS) is calculated, and it is found that the initial relative translational energy and vibration excitation have a significant effect on ECSs, while rotational excitation has little influence; the rate coefficient of the high rovibrational state of O2 molecules at high temperatures is studied, and it is found that when the vibrational level ν of O2 is in the range of 0-15, the value of log10k(T, ν, j) with the vibrational level ν is almost linear, while when ν is greater than 15, it becomes gentle with the increase in ν. Finally, the state-to-state rate coefficients are calculated; our results supply the advantageous state-to-state process data in the NO2 system, and they are useful for further studying the related hypersonic gas flow at very high temperature.