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The purpose of this paper is to study the effects of boundary‐layers bleeding on performance parameters of hypersonic inlets.

The inner flowfield of a hypersonic inlet at different bleeding rates is simulated with a Reynolds‐averaged Navier‐Stokes solver using a renormalization group k‐ε turbulence model.

In contrast with no bleeding, the performance parameter of hypersonic inlets without backpressure is reduced slightly, but the flow uniformity is improved. The interaction between boundary layers and shocks is weakened at the action of the bleeding, which leads to that the boundary‐layers separations at the entrance of the isolator caused by the high‐backpressure occur later, and it can improve the maximum backpressure ratio of hypersonic inlets. With the bleeding rate increasing, the maximum backpressure ratio of hypersonic inlets is added, while the total‐pressure recovery coefficient and mass‐captured coefficient are reduced.

This paper is a useful reference to the design and performance improvement of hypersonic inlets and propulsion systems.

This paper aims to present a novel structural reliability analysis scheme with considering the structural strength degradation for the wing spar of a generic hypersonic aircraft to guarantee flight safety and structural reliability.

A logarithmic model with strength degradation for the wing spar is constructed, and a reliability model of the wing spar is established based on stress-strength interference theory and total probability theorem.

It is demonstrated that the proposed reliability analysis scheme can obtain more accurate structural reliability and failure results for the wing spar, and the strength degradation cannot be neglected. Furthermore, the obtained results will provide an important reference for the structural safety of hypersonic aircraft.

The proposed reliability analysis scheme has not implemented in actual flight, as all the simulations are conducted according to the actual experiment data.

The proposed reliability analysis scheme can solve the structural life problem of the wing spar for hypersonic aircraft and meet engineering practice requirements, and it also provides an important reference to guarantee the flight safety and structural reliability for hypersonic aircraft.

To describe the damage evolution more accurately, with consideration of strength degradation, flight dynamics and material characteristics of the hypersonic aircraft, the stress-strength interference method is first applied to analyze the structural reliability of the wing spar for the hypersonic aircraft. The proposed analysis scheme is implemented on the dynamic model of the hypersonic aircraft, and the simulation demonstrates that a more reasonable reliability result can be achieved.

This paper aims to present an aerothermodynamic analysis of a new concept of a small hypersonic airplane. Aerodynamics characteristics for different flow conditions encountered during the missions are analyzed. The effects of elevons deflection for pitch control and of the presence of engines on aerodynamic performances are also investigated for different flight conditions. The effects of boundary layer laminar–turbulent transition on aerodynamic heating are studied to preliminarily identify proper materials that can sustain the hypersonic phase.

Aerodynamic characteristics are predicted by means of the semi-empirical aerodynamic prediction code Missile DATCOM and computational fluid dynamics simulations. Computational fluid dynamics analysis is also performed to investigate aerodynamic heating phenomenon.

Major discrepancies between the results offered by the two methods have been registered in transonic regime, whereas in subsonic and super-hypersonic conditions, Missile DATCOM confirms to be a suitable tool for preliminary design steps. The results of the analysis show that for the identification of the materials that can sustain the hypersonic phase, the turbulent solution must be taken into account. Carbon fiber reinforced ceramics composite materials seem particularly well suited for the nose, wing and vertical tail leasing edges and control surfaces, while titanium alloys could be used for the rest of the vehicle surface.

This new concept of vehicle is designed both for point-to-point medium range hypersonic transportation and long duration suborbital space tourism missions, by integrating available technologies developed for aeronautical and space systems.

The purpose of the paper is to design a robust control system for a generic hypersonic vehicle which includes dynamic nonlinear, open loop unstable and parametric uncertainties.

For a complex longitudinal model of a generic hypersonic vehicle which includes dynamic nonlinear, open loop unstable and parametric uncertainties, a nonlinear dynamic inverse (NDI) approach combined with proportional differential (PD) control is used to design a strong robust control system to deal with the sensitivity to changes of atmosphere condition. In this way, a simple genetic algorithm is used to search a group of parameters of the control system to satisfy the specific performance indices. Then parametric uncertainties are considered to verify the robustness of the control system.

The PD hypersonic vehicle control system using NDI approach can satisfy the specific flight performance. And it has strong robustness under the parametric uncertainties.

The paper fulfills a complete process of the nonlinear control system design for a generic hypersonic vehicle. And, the simulation results show the efficiency and robustness of the control system.

As to different angles of attack and nonlinear problems caused by high temperatures in coexisting hypersonic aircraft, people mainly rely on fluid software for research but lack analysis of flow mechanisms. Owing to computational difficulties, few people use numerical algorithms to combine them for discussion. Hence, this study aims to make a deep inquiry into the laminar flow and heat transfer of compressible Newtonian fluid in hypersonic aircraft with small attack angles.

In this paper, on the basis of mass, momentum and energy conservation laws, the governing equations of the hypersonic boundary layer are established. Viscosity, specific heat capacity and thermal conductivity are considered nonlinear functions concerning temperature. In virtue of the MacCormack finite difference method, the stationary numerical solutions are solved directly, and the validity of the algorithm is verified.

The results demonstrate that at Mach number 5, compared to the 0° attack angle, the maximum temperature near-wall at the 3° attack angle increases by about 25%. An enjoyable phenomenon is discovered, where the position corresponding to the maximum wall shear force shifts back as the attack angle and Mach number increase. The relationship between the near-wall maximum temperature versus attack angle and Mach number is fitted through numerical calculation results.

Empirical formulas can be used to estimate heat transfer characteristics at small attack angles, which will guide the design of aircraft thermal protection systems.

Waverider has high lift to drag ratio and will be an idea aerodynamic configuration for hypersonic vehicles. But a structure permitting aerodynamic like waverider is still difficult to generate under airframe’s geometric constrains using traditional waverider design methods. And furthermore, traditional waverider’s aerodynamic compression ability cannot be easily adjusted to satisfy the inlet entrance requirements for hypersonic air-breathing vehicles. The purpose of this paper is to present a new method named osculating general curved cone (OCC) method aimed to improve the shortcomings of traditional waveriders.

A basic curved cone is, first, designed by the method of characteristics. Then the waverider’s inlet captured curve and front captured tube are defined in the waverider’s exit plane. Osculating planes are generated along the inlet captured curve and the designed curved cone is transformed to the osculating planes. Streamlines are traced in the transformed curved cone flow field. Combining all streamlines which have been obtained, OCC waverider’s compression surface is generated. Waverider’s upper surface uses the free stream surface.

It is found that OCC waverider has good volumetric characteristics and good flow compression abilities compared with the traditional osculating cone (OC) waverider. The volume of OCC waverider is 25 per cent larger than OC waverider at the same design condition. Furthermore, OCC waverider can compress incoming flow to required flow conditions with high total pressure recovery in the waverider’s exit plane. The flow uniformity in the waverider exit plane is quite well.

The analyzed results show that the OCC waverider can be a practical high performance airframe/forebody for hypersonic vehicles. Furthermore, this novel waverider design method can be used to design a structure permitting aerodynamic like waverider for a practical hypersonic vehicle.

The paper puts forward a novel waverider design method which can improve the waverider’s volumetric characteristics and compression abilities compared with the traditional waverider design methods. This novel design approach can extend the waverider’s applications for designing hypersonic vehicles.

The Fay and Riddell (F–R) formula is an empirical equation for estimating the stagnation-point heat flux on noncatalytic and fully catalytic surfaces, based on an assumption of equilibrium. Because of its simplicity, the F–R has been used extensively for reentry flight design as well as ground test facility applications. This study aims to investigate the uncertainties of the F-R formula by considering velocity gradient, chemical species at the boundary layer edge, and the thermochemical nonequilibrium (NEQ) behind the shock layer under various hypersonic NEQ flow environments.

The stagnation-point heat flux calculated with the F–R formula was evaluated by comparison with thermochemical NEQ calculations and existing flight experimental values.

The comparisons showed that the F–R underestimated the noncatalytic heat flux, because of the chemical composition at the surface. However, for fully catalytic heat flux, the F–R results were similar to values of surface heat flux from thermochemical NEQ calculations, because the F–R formula overestimates the diffusive heat flux. When compared with the surface heat flux results obtained from flight experimental data, the F–R overestimated the fully catalytic heat flux. The error was 50% at most.

The results provided guidelines for the F–R calculations under hypersonic flight conditions and for determining the approximate error range for noncatalytic and fully catalytic surfaces.

The purpose of this paper is to explore the characteristics of hypersonic flow past a blunt body.

The implicit finite volume schemes are derived from axisymmetric Navier–Stokes equations by means of AUSM+ and LU-SGS methods, and programmed in FORTRAN. Based on the verified result that a 2D axisymmetric chemical equilibrium flow has a good agreement with the literature, the characteristics of hypersonic flow past a sphere are simulated by using four different models which involve four factors, namely, viscous, inviscid, equilibrium and calorically perfect gas.

Compared with the calorically perfect gas under hypervelocity condition, the shock wave of the equilibrium gas is more close to the blunt body, gas density and pressure become bigger, but gas temperature is lower due to the effect of real gas. Viscous effects are not obvious in the calculations of the equilibrium gas or the calorically perfect gas. In a word, the model of equilibrium gas is more suitable for hypersonic flow and the calculation of viscous flow has a smaller error.

The computer codes are developed to simulate the characteristics of hypersonic flows, and this study will be helpful for the design of the thermal protection system in hypersonic vehicles.

The purpose of this paper is to propose a new approach for aerodynamic parameter identification of hypersonic vehicles, which is based on Pigeon-inspired optimization (PIO) algorithm, with the objective of overcoming the disadvantages of traditional methods based on gradient such as New Raphson method, especially in noisy environment.

The model of hypersonic vehicles and PIO algorithm is established for aerodynamic parameter identification. Using the idea, identification problem will be converted into the optimization problem.

A new swarm optimization method, PIO algorithm is applied in this identification process. Experimental results demonstrated the robustness and effectiveness of the proposed method: it can guarantee accurate identification results in noisy environment without fussy calculation of sensitivity.

The new method developed in this paper can be easily applied to solve complex optimization problems when some traditional method is failed, and can afford the accurate hypersonic parameter for control rate design of hypersonic vehicles.

In this paper, the authors converted this identification problem into the optimization problem using the new swarm optimization method – PIO. This new approach is proved to be reasonable through simulation.

THE purpose of this paper is to review the major structural design problems common to all hypersonic air vehicles and describe some of the current research results to indicate the structures and materials technology produced during the initial phase of a new era in aeronautics. To further limit the scope of the discussion and to focus it on a particular application, these structural problems will be discussed in relation to a hypersonic commercial air transport that might be ready for intercontinental airline service 15 to 20 years' hence.