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The purpose of this paper is to optimize the downrange for hypersonic boost‐glide (HBG) missile under near‐real condition, and to validate the suitability of proposed wall cooling materials.

The trajectory optimization problem is characterized by a boost phase followed by a glide phase. A multi‐phase trajectory optimization tool is adopted to optimize the downrange. The associated optimal control problem has been solved by selecting a direct shooting method. The dynamics has been transcribed to a set of nonlinear constraints and the arising nonlinear programming problem has been solved through a sequential quadratic programming solver. An aerothermodynamics analysis method is introduced to calculate the aerodynamic heating at nose, leading edge, and ventral centerline regions.

HBG missile is suitable for long‐range attack, and the optimal trajectory solved is a novel boost‐glide‐skip trajectory, i.e. boost firstly, glide secondly, and skip at last. The proposed wall materials are valid.

This paper provides further study on the methods of trajectory design and aerothermodynamics analysis for HBG missile.

The purpose of this paper is to attempt an aerospaceplane design with the objective of Low-Earth-Orbit-and-Return-to-Earth (LEOARTE) under the constraints of safety, low cost, reliability, low maintenance, aircraft-like operation and environmental compatibility. Along the same lines, a “sister” point-to-point flight on Earth Suborbital Aerospaceplane is proposed.

The LEOARTE aerospaceplane is based on a simple design, proven low risk technology, a small payload, an aerodynamic solution to re-entry heating, the high-speed phase of the outgoing flight taking place outside the atmosphere, a propulsion system comprising turbojet and rocket engines, an Air Collection and Enrichment System (ACES) and an appropriate mission profile.

It was found that a LEOARTE aerospaceplane design subject to the specified constraints with a cost as low as 950 United States Dollars (US$) per kilogram into Low Earth Orbit (LEO) might be feasible. As indicated by a case study, a LEOARTE aerospaceplane could lead, among other activities in space, to economically viable Space-Based Solar Power (SBSP). Its “sister” Suborbital aerospaceplane design could provide high-speed, point-to-point flights on the Earth.

The proposed LEOARTE aerospaceplane design renders space exploitation affordable and is much safer than ever before.

This paper provides an alternative approach to aerospaceplane design as a result of a new aerodynamically oriented Thermal Protection System (TPS) and a, perhaps, improved ACES. This approach might initiate widespread exploitation of space and offer a solution to the high-speed “air” transportation issue.

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 this paper is to present the more accurate estimation of aero-heating for the ablative 3-D noses by using the viscous shock layers and similarity of viscous boundary layer methods.

The combination of viscous shock layer, similarity of viscous boundary layer (SVBL) methods, Park ablation and Baldwin–Lomax turbulent models is presented in this paper. The proposed method reduces computational memory and run time as compared to the time marching algorithms during flight trajectory. Therefore, the space marching algorithm and finite difference method is used, and the governing equations are transferred into curvature coordinate by using the mapping terms.

The solving for an ogive nose during flight trajectory shows that the convergence of this technique is fast as compared to the user defined function based on the fluent solvers, program to axisymmetric regular geometry code and other research. The results of this research are validated by the mentioned research studies. The relative error for the aero-heating, species concentration of the shock layer gas mixture because of dissociation/ionization of air and surface ablation results is less than 6, 5 and 11 per cent, respectively.

The required time for an aerodynamic design of hypersonic noses reduces as the induced aero-heating is one of the principal design parameters in standpoint aerodynamic, structural and other terms. The magnitude of this parameter, surface temperature and surface recess because of ablation should be corrected during flight trajectory.

The results of this research are applicable for aerospace industries.

The originality of this paper is 90 per cent.

In this paper it is described how trajectory data is used to calculate the distribution of temperature throughout the radome of a missile. The problem divides itself into two parts, one being the solution of the heat conduction equation within the medium of the radome, the second being the solution of the boundary condition at the surface. The method is easily applied to many other problems where a medium is heated by forced convection and in fact three independent problems have been solved using the method described here.

The transpiration has been recognized as one of the most effective thermal protection methods for future hypersonic vehicles. To improve efficiency and safety, it is urgent to optimize the design of the transpiration system for heat and drag reduction. The purpose of this paper is to investigate the effects of transpiration on heat and drag reduction.

A chemical nonequilibrium flow model with the transpiration is established by using Navier–Stokes equations, the shear-stress transport turbulence model, thermodynamic properties and the Gupta chemical kinetics model. The solver programmed for this model is verified by comparing with experimental results in the literature. Effects of air injection on the flow field, the aerodynamic resistance and the surface heat flux are calculated with the hypersonic flow past a blunt body. Furthermore, a modified blocking coefficient formula is proposed.

Numerical results show that the transpiration can reduce the aerodynamic resistance and the surface heat flux observably and increase the shock wave standoff distance slightly. It is also manifested that the modified formula is in better agreement with the wind tunnel test results than the original formula.

The modified formula can expand the application range of the engineering method for the blocking coefficient. This study will be beneficial to carry out the optimal design of the transpiration system.

The aim of this study is to design a guidance method to generate a smoother and feasible gliding reentry trajectory, a highly constrained problem by formalizing the control variables profile.

A novel accelerated fractional-order particle swarm optimization (FAPSO) method is proposed for velocity updates to design the guidance method for gliding reentry flight vehicles with fixed final energy.

By using the common aero vehicle as a test case for the simulation purpose, it is found that during the initial phase of the longitudinal guidance, there are oscillations in the state parameters which cause to violate the path constraints. For the glide phase of the longitudinal guidance, the path constraints have higher values because of the increase in the atmosphere density.

The violation in the path constraints may compromise the flight vehicle safety, whereas the enforcement assures the flight safety by flying it within the reentry corridor.

An oscillation suppression scheme is proposed by using the FAPSO method during the initial phase of the reentry flight, which smooths the trajectory and enforces the path constraints partially. To enforce the path constraints strictly in the glide phase, ultimately, another scheme by using the FAPSO method is proposed. The simulation results show that the proposed algorithm is efficient to achieve better convergence and accuracy for nominal as well as dispersed conditions.

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.

This paper aims to present the problem of heating the damaged insulation of an orbiter.

Changes of the insulation’s thermal properties, made by adding conductive material of high value of specific heat in a form of a dope to the protective layer, were examined. An iterative algorithm determining a variable of dope concentration in the material was developed.

Determination of distribution of conductive material concentration was made for materials which, after verification, demonstrated the most beneficial effect on protective properties of the modified insulation layer. The problem of determining the distribution of metal filings concentration in the insulation layer of the coating belongs to inverse heat conduction problems.

Change of properties was to enable time extension of the LI900 insulation tile heating up to the maximal temperature and, additionally, to lowering this temperature.

In order to improve the engine reliability and efficiency, an effective way is to reform the turbine blade tip conformation. The paper aims to discuss this issue.

The present research provides several novel tip-shaping structures, which are considered to control the blade tip loss. Four different tip geometries have been studied: flat tip, squealer tip, flat tip with streamwise ribs and squealer tip with streamwise ribs. The tip heat transfer and leakage flow are both analyzed in detail, for example the tip heat transfer coefficient, tip flow and local pressure distributions.

The results show that the squealer seal and streamwise rib can reduce the tip heat transfer and leakage loss, especially for the squealer tip with streamwise ribs. The tip and near-tip flow patterns at the different locations of axial chord reflect that both the squealer seal and streamwise rib structure can control the tip leakage flow loss. In addition, the analysis of the aerodynamic parameters (the static pressure and turbine efficiency) also indicates that the squealer tip with streamwise ribs obtains the highest adiabatic efficiency with an increase of 2.34 percent, compared with that of the flat tip case.

The analysis of aerothermal and dynamic performance can provide a reference for the blade tip design and treatment.