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This study reviews the application of the C/SiC composites in the thermal protection systems (TPS). Special attention is put on the structure design and optimization of the C/SiC shell panel. The comparisons of several TPS show that the ceramic matrix composite (CMC) is one of the most potential materials for TPS. The crucial role of multi‐disciplinary design and optimization is pointed out in the future design of C/SiC thermal protection structure. This review provides a guideline on the current and future study of C/SiC thermal protection structure.

Common Aero Vehicles (CAVs) are relatively small‐size, un‐powered, self‐maneuvering vehicles equipped with a variety of weapons and launched from space. One of the major obstacles hampering a full the realization of the CAV concept is a present lack of lightweight, high‐temperature insulation materials which can be used for construction of the CAV’s thermal protection system (TPS). A computational analysis is utilized in the present work to examine the suitability of a carbon‐based, coal‐derived foam for the TPS applications in the CAVs. Toward that end, a model is developed for the high‐temperature effective thermal conductivity of foam‐like materials. In addition, an insulation sizing procedure is devised to determine the minimum insulation thickness needed for thermal protection of the vehicle structure at different sections of a CAV. It is found that the carbon‐based foam material in question can be considered as a suitable TPS insulation material at the leeward side and at selected portions of the windward side of a CAV (specifically the portions which are further away from the vehicle nose).

For a thermal protection system (TPS) of long endurance hypersonic flight vehicle (HFV), its thermal insulation property not only determines by the manufactured morphology but also changes along time. A thermal conductivity prediction model for aerogel considering heat treatment effect is carried out and applied to solve the heat conduction problem of a TPS. The aim of this study is to provide theoretical and numerical references for further development of aerogels applying to TPSs.

A thermal conductivity prediction model for aerogel is established considering treatment effect. The heat conduction problem of a TPS is derived and solved by combining the differential quadrature method and the Runge–Kutta method. The prediction results of aerogel thermal conductivities are verified by comparing with those in literature, while the calculated temperature field of TPS is verified by comparing with that by ABAQUS.

Numerical results show that when applying the current prediction model, the calculated high temperature area in the aerogel layer is narrowed due to the decrease of the thermal conductivity during heat treatment process.

This study will be beneficial to carry out the precise design of TPS for long endurance HFVs.

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.

A multi-disciplinary robust design optimization method for micro Mars entry probe (no more than 0.8 m in diameter) is proposed. The purpose of this paper is to design a Mars entry probe, not only the geometric configuration, but the trajectory and thermal protection system (TPS). In the design optimization, the uncertainties of atmospheric and aerodynamic parameters are taken into account. The probability distribution information of the uncertainties are supposed to be unknown in the design. To ensure accuracy levels, time-consuming numerical models are coupled in the optimization. Multi-fidelity approach is designed for model management to balance the computational cost and accuracy.

Uncertainties which cannot defined by usual Gaussian probability distribution are modeled with degree of belief, and optimized through with multiple-objective optimization method. The optimization objectives are set to be the thermal performance of the probe TPS and the corresponding belief values. Robust Pareto front is obtained by an improved multi-objective density estimator algorithm. Multi-fidelity management is performed with an Artificial Neural Network (ANN) surrogate model. Analytical model is used first, and then with the improvement of accuracy, rather complex numerical models are activated. ANN updates the database during the optimization, and makes the solutions finally converge to a high-level accuracy.

The optimization method provides a way for conducting complex design optimization involving multi-discipline and multi-fidelity models. Uncertainty effects are analyzed and optimized through multi-disciplinary robust design. Because of the micro size, and uncertain impacts of aerodynamic and atmospheric parameters, simulation results show the performance trade-off by the uncertainties. Therefore an effective robust design is necessary for micro entry probe, particularly when details of model parameter are not available.

The optimization is performed through a new developed multi-objective density estimator algorithm. Affinity propagation algorithm partitions adaptively the samples by passing and analyzing messages between data points. Local principle component techniques are employed to resample and reproduce new individuals in each cluster. A strategy similar to NSGA-II selects data with better performance, and converges to the Pareto front. Models with different fidelity levels are incorporated in the multi-disciplinary design via ANN surrogate model. Database of aerodynamic coefficients is updated in an online manner. The computational time is greatly reduced while keeping nearly the same accuracy level of high fidelity model.

This paper aims to address a significant issue related to the coupled and uncoupled treatment of the thermal and dynamic problems in the optimization of aeroassisted orbital maneuvers and the simultaneous optimal sizing of the associated heat shields. The literature generally focuses on decoupled treatments that reduce the computational load; in this manner, consequently, a decrease in the representativity of the solution manifests. The general operating mode first optimizes the trajectory and subsequently defines the optimal heat shield design based on that trajectory.

This paper analyzes the impact of both treatments on the evaluation of the convenience of an aeroassisted maneuver with respect to an equivalent purely propulsive exoatmospheric maneuver in relation to the achievable total mass savings of the propellant and the heat shield. Two case studies are analyzed via an optimization methodology that references genetic algorithms: the first case study is related to an aerobraking maneuver and the second case study is related to an orbital plane change.

The results demonstrate that the adoption of decoupling produces conservative solutions, i.e. unfavorable estimates, with a lower level of convenience of the aeroassisted technique compared to equivalent purely propulsive exoatmospheric maneuvers.

This type of analysis can provide an appropriate discernment criterion for the selection of the modus operandi based on the available computational power and the desired level of representativity.

Silicon carbide multilayer composites containing short carbon fibers (C_sf/SiC) were prepared by tape casting and pressureless sintering. The C fibers were dispersed with dispersants into a solvent mixture firstly and then mixed with SiC slurry to make green C^sf/SiC tapes. Triton X100 was found to be the best dispersant for Toho Tenax HTC124 fibers. Fibers resulted homogeneously distributed in the tape and tended to align fairly well along the tape casting direction. The addition of short C fibers hindered the shrinkage in the plane containing the fibers during sintering. The resulting microstructure of the composite materials was investigated. This kind of composite layers could be integrated in a thermal protection system (TPS) structure, since the outer dense SiC layers can provide excellent oxidation resistance and good heat conductivity in the plane, while C_sf/SiC layers in the middle of the multilayer architecture could grant low thermal conductivity through the TPS thickness.

The effective heat capacity is a key index to estimate the thermal protection performance of charring ablative materials in reentry vehicles subjected to aerodynamic heat loads. The purpose of this paper is to investigate the effects of gradient density on the effective heat capacity.

Based on the Fourier law and the pyrolysis interface model, the authors establish the governing equations for the transient heat conduction with variable density, and then simulate one-dimensional transient thermal behavior of a homogeneous and three types of non-homogeneous charring ablative material in reentry capsules by using the implicit numerical method.

The moving rate of pyrolysis interface and the surface temperature of charring ablative material depend on not only the surface heating history, but also the gradient density. And the gradient density can improve the insulation performance of charring materials, e.g. the effective heat capacity in the bilinear design is larger than that in the homogeneous design under a given heat flux condition.

This study will help the design of the thermal protection system in reentry vehicles.

The purpose of the paper is to study the variation of optimal burnout angle at the end of the ascent phase and the optimal control deflection during the glide phase, that would maximize the downrange performance of a hypersonic boost-glide waverider, with variation in heat rate and integrated heat load limit.

The approach used is to model the boost phase so as to optimize the burnout conditions. The nonlinear, multiphase, constraint optimal control problem is solved using an hp-adaptive pseudospectral method.

The constraint heat load results for the waverider configuration reveal that the integrated heat load can be reduced by more than half with only 10 per cent penalty in the overall downrange of the hypersonic boost-glide vehicle, within a burnout speed range of 3.7 to 4.3 km/s. The angle-of-attack trim control requirements increase with stringent heat rate and integrated heat load bounds. The normal acceleration remains within limits.

The trajectory results imply lower thermal protection system weight because of reduced heat load trajectory profile and therefore lower thermal protection system cost.

The research provides further study on the trajectory design to the hypersonic boost-glide vehicles for medium range application.

Two decades ago, the space age dawned, and we were awakened to a realm of technological possibilities beyond any we had imagined. Since that time, we have linked continents with communication satellites, sent probes to other planets, and seen men on the moon. And, although we have made achievements that were almost inconceivable 30 years ago, we are still far from realising our potential for a future in space.