Thermal Protection Design Research Articles

Three-dimensional pore-scale study of the directional heat transfer in a high thermal conductivity carbon/carbon composite protection system

High-efficiency thermal protection is becoming a critical design criterion for supersonic aircraft. Currently, there is significant research on accurately and efficiently designing heat protection systems, particularly at the high temperature stagnation point region of supersonic aircraft. In this work, a three-dimensional high thermal conductivity carbon/carbon composite thermal protection system with a directional heat transfer function is proposed to realize a high-efficiency thermal protection design in the stagnation point region. A comprehensive method that combines the lattice Boltzmann method (LBM) and finite volume method (FVM) was designed to investigate the heat transfer process in the proposed structure at the pore scale. The FVM was applied to calculate the heat radiative information, which was required to solve the energy equation with the LBM. A failure temperature is defined to quantitatively describe how well the heat transfer occurs in the designed direction in the proposed structure. The effects of the temperature of the stagnation point, thermal conductivity of the carbon fibers, carbon fiber diameters, and position of the carbon fibers on the heat transfer in the proposed structure were studied in detail. The results show that the proposed comprehensive method can predict the heat transfer in the proposed structure accurately. A competitive relationship exists between the directional heat conduction and the radiation heat transfer. The radiation heat transfer became more dominating as the temperature increased. The failure temperature for the directional heat transfer in the proposed structure increased with the thermal conductivity of the carbon fibers and the carbon fiber diameters, while it decreased as the distance between the position of the carbon fibers and the nose of the blunt structure increased. The above phenomena will be helpful for designing high-performance thermal protection systems with carbon/carbon composites.

Aerothermodynamic study of Two-Stage-To-Orbit system composed of wide-speed-range vehicle and rocket

Hypersonic vehicles with Mach numbers greater than five exhibit complex shock wave interactions, accompanied by locally enhanced heat flux near the shock-affected surface. To investigate the unclear interstage thermal environment in a Two-Stage-To-Orbit system, a model composed of a wide-speed-range vehicle and a reusable rocket is employed in an aerothermodynamic study of the Two-Stage-To-Orbit system considering the interstage interactions performed at freestream Mach number = 6. The innovations of this paper are the complex model, which is highly similar to the real situation, and the study of the thermal environment between stages near different reflection positions, whereby the shock wave interference that causes different forms of heat flux is divided into two categories. The thermal design of the Two-Stage-To-Orbit system should especially consider the head of the wide-speed-range vehicle, interstage distance, and the swept angle of canted fins and wings, as well as their width. This paper adopts the second-order AUSMDV (a variant of the Advection Upstream Splitting Method) scheme for spatial discretization and Lower–Upper Symmetric Gauss–Seidel method for time discretization in formulating the Reynolds-averaged Navier–Stokes equations, which are then solved by the finite volume method and Menter’s shear stress transport k-ω turbulence model. Numerical analysis sheds light on the complex shock structure and the reasons for the considerable increase in wall temperature and aerothermal loads at critical parts of various reflected shock positions. The incident shock wave generated by the head apex of the rocket constantly reflects and gradually weakens between the two stages, which brings about shock wave/boundary-layer interaction and gives rise to elevated heat flux. The moderate incident shock and the weak reflected shock are primarily responsible for the lack of boundary-layer separation, although the expansion caused by the convex curvature of the head of the wide-speed-range vehicle also contributes. A non-reflected part of the incident shock wave interacts with the windward surface of the canted fins and wings, producing shock wave/boundary-layer interaction and a V-shaped heat flux distribution that extends to the leading edges of the wings. As the angle of attack decreases, the intensity of the incident shock and the reflected shock gradually weakens, leading to reduced heat flux on the windward surface of the wide-speed-range vehicle. The range over which the incident shock impinges the walls also narrows, corresponding to the smaller heat flux distribution. These aerothermal studies comprehensively determine the law of thermal changes, which is significant for the design of thermal protection in Two-Stage-To-Orbit systems.

Acoustic-wave-induced cooling in onset of hypersonic turbulence

We report a newly identified aerodynamic cooling mechanism in the onset of hypersonic wall-bounded turbulence. We first experimentally investigated a flared cone with a smooth surface in a Ma 6 wind tunnel using fast-response pressure sensors, Rayleigh scattering flow visualization, and infrared thermography, which confirmed a cooled region (denoted as CS) downstream of a highly heated region (denoted as HS) on the model, as shown by Franko and Lele [J. Fluid Mech. 730, 491–532 (2013)] and Sivasubramanian and Fasel [J. Fluid Mech. 768, 175–218 (2015)]. We then performed calculations based on both linear stability theory and direct numerical simulations to understand this mechanism. We found that the phase difference ϕpθ between the periodic pressure and dilatation waves plays an important role in the interchange between thermal and mechanical energy in a hypersonic wall-bounded flow. Using porous steel to modify the model surface’s sound admittance, we experimentally show that it is possible to modify the cosine value of ϕpθ to be negative near the wall and thus reduce the temperature growth. These results can provide insight into the thermal protection design of future hypersonic vehicles.

Design of thermal protection based on open cell carbon foam structure optimization

Open cell carbon foams with porosity ranging between 82 and 98% can be used as efficient thermal insulation in many high-temperature applications including solar and planetary probes, whose structures and systems are exposed to extreme heat loads. Physical properties of cellular materials are determined not only by thermal and optical properties of solid phase but also significantly by their morphology and fabrication technology. This implies the possibility to create open-cell materials with desirable properties, optimal for specific operating conditions. The paper presents a methodology for optimal design of multilayer thermal protection based on high porosity open cell carbon foam, taking into account the carbon foam morphology. The traditional thermal design problem statement implies the determination of layers thickness for multi-layer thermal insulation, ensuring required operational temperature on the boundaries of layers and minimum of total mass of system. In this work the cell diameter and porosity, that characterize the material’s morphology, are chosen along with thickness of layers in order to obtain an additional weight advantage of thermal insulation. The optimization problem is solved using the algorithm based on the projected Lagrangian method with the quadratic subproblem. To illustrate the implementation of the developed algorithm and the corresponding software, the problem of choosing of the optimal layer thickness for the multilayer heat shield of the solar probe along with the cell diameter and porosity of carbon foam, forming one of the layers, is considered.

Numerical Prediction of Flow Field and Aerodynamic Heating in the Gap of Hypersonic Vehicle

Prediction of flow field and thermal environment in the gap is crucial to the design of hypersonic vehicle thermal protection system. The compressible Navier-Stokes equation is solved to study the flow field and aerodynamic heating in the gap. The effect of inflow parameter on the aerodynamic heating is analyzed. The results show that there are vortex motions in the gap. The heat transfer in the gap mainly depends on the convection in the top region. The heat flux along the gap wall distributes in U-shape, and at the outlet of the gap is highest, which is the key point in the thermal protection design. As the angle of attack increases, the number of vortices will decrease gradually, and wall heat flux will increase, especially at the gap exit. The research results provide references for the design and optimization of thermal protection system.

Методы управления инвестиционными энергосберегающими проектами

The article presents the study results of organizational and economic aspects of managing energy-saving investment projects. The requirements for the energy service construction contract are specified. The authors propose an economic-mathematical method for determining the optimal design of thermal protection of buildings taking into account the law of diminishing marginal utility.

On the Universal Law of the Decomposition of Thermal-Protective Binder Composite Materials at High Temperatures

Based on the known (rated) values of the densities and temperatures at the beginning and end of the decomposition of thermal-protective binder composite materials, as well as on the exponential nature of the decomposition (pyrolysis) of the binder composite materials, the law of binder pyrolysis is established with the formation of a gas component and a porous coke residue, through which the pyrolysis gases are filtered and are blown into the high-temperature boundary layer. The law is universal in nature and suitable for almost any heat-shielding composite materials, since it does not use the chemical kinetics of decomposition, which are difficult to formalize. The use of this decomposition law will greatly simplify the mathematical modeling of complex processes for the design of thermal protection for high-speed and hypersonic aircraft. An analysis of the obtained calculation results is presented.

An optimal design of thermal protection based on materials morphology

The paper presents a methodology for the optimal design of multilayer thermal protection based on high-porosity open-cell foam, taking into account the foam morphology. The vector of design parameters includes the thicknesses of layers, porosity and cell diameter of open-cell foam and should ensure required operation temperature on the boundaries of layers and a minimum of the total mass of the system. The optimization problem is solved using a computational scheme, which combines the projected Lagrangian method with the quadratic subproblem and the penalty function method. The penalty function method provides a good initial estimate of the optimal parameters’ values for the projected Lagrangian method with excellent local convergence properties. To illustrate the implementation of the developed algorithm and the corresponding software, the problem of choosing the optimal layer thicknesses for the multilayer thermal protection of a spacecraft along with the cell diameter and porosity of foam is considered.

Assessing Turbulence Models on the Simulation of Launch Vehicle Base Heating

Launch vehicles suffer from severe base heating during ascents. To predict launch vehicle base heat flux, the computational fluid dynamics (CFD) tools are widely used. The selection of the turbulence model determines the numerical simulation results of launch vehicle base heating, which may instruct the thermal protection design for the launch vehicle base. To assess performances, several Reynolds-averaged turbulence models have been investigated for the base heating simulation based on a four-nozzle launch vehicle model. The finite-rate chemistry model was used for afterburning. The results showed that all the turbulence models have provided nearly identical mean flow properties at the nozzle exit. Menter’s baseline (BSL) and shear stress transport (SST) models have estimated the highest collision pressure and have best predicted base heat flux compared to the experiment. The Spalart-Allmaras (SA) model and the renormalization group (RNG) model have performed best in temperature estimation, respectively, in around r/rb=0~0.2 and r/rb=0.6~1. The realizable k‐ε (RKE) model has underestimated the reverse flow and failed to correctly reflect the recirculation in the base region, thus poorly predicted base heating. Among all the investigated turbulence models, the BSL and SST models are more suitable for launch vehicle base heating simulation.

Numerical study on the effect of rocket plume to the jet flow deflector in liquid rocket engine test stage

Jet flow deflector is one of the most important facility in liquid rocket engine test stage. It is necessary to investigate the effect of rocket plume to the jet flow deflector. Based on the theory of computational fluid dynamics, the flow field of a rocket engine plume is simulated and analysed by numerical method. The velocity field and temperature field of the plume in rocket engine test are obtained. The structural characteristics of engine plume flow field and the impact effect on flow deflector are analysed. Finally, results are applied to the thermal protection design of the test stage.

Characteristics of the Mechanism of Thermal Power Destruction of Carbon Materials in a Supersonic High-Temperature Air Flow

An experimental simulation is performed for the mechanism of the destruction of carbon heat-shielding materials (polycrystalline graphite and carbon composites) with a rough surface passed by a supersonic high-temperature air flow. At high velocities of the oncoming flow, the roughness of such materials may correspond to the thickness of the boundary layer. Such properties of the surface contribute to a more intensive destruction of materials due to the mechanical carryover of mass caused by the gas dynamics of the flow and the thermochemical processes on a rough surface. This paper analyzes the methods and means of research under the ground conditions of the mechanism of the destruction of carbon heat-shielding materials that can be used in the design of thermal protection for rocket-space and high-speed aviation equipment.

Hydrogen flow and heat transfer characteristic analysis in cooling channel wall with the spherical convexity structure

In the thermal protection design for rocket engines chamber, the internal wall always suffers from high combustion temperature and small cooling channel size, thus an efficient method by changing inner wall structure in cooling channel is investigated. This paper proposed a structure with spherical convexity on the inner wall for improving the entire cooling performance, weakening thermal stratification phenomenon, and protecting the overheated structure. Optimized spacing is obtained through comparison to achieve better cooling effect and smaller flow pressure drop. An enhanced disturbance, which stems from the newly introduced spherical convexity surface to control coolant flow, finally improves the convective heat transfer efficiency. The results show that the spherical convexity structure of a depth 0.2 mm and spacing 2.5 mm is proper for enhancing heat transfer and weakening thermal stratification phenomenon in the channel with cross-section 2 mm*2 mm. Compared to the smooth channel, the (Nu/Nu0)/(f/f0)1/3 is increased by 55%, and the maximum temperature is decreased by 10% but the pressure drop is only increased by a maximum 14.5%.

INVESTIGATION OF THE INFLUENCE OF THERMAL AND CLIMATE EFFECTS ON THE PERFORMANCE OF TILED THERMAL PROTECTION OF SPACECRAFT

The article discusses the thermal protection of spacecraft under various types of thermal and climatic influences. During the pre-launch preparation, several types of icing are possible on the surface of the tiled structure, consisting of quartz heat-shielding tiles. It is very important in the process of operation, to determine normal and shear stresses during icing and the impact of ice formed in the tile gaps. It is necessary to calculate the strength and performance of the tiled thermal protection in active stages of flight of spacecraft. A mathematical model was constructed that allows, with a sufficient degree of accuracy, to research the tiled structure for strength, shear and tearing of the tile when exposed to ice. Calculations for strength, shear, and tearing of tiles under the influence of ice were performed. Theoretical and experimental studies were compared. The performed calculations made it possible to show that the tile will not collapse under the considered load. The strength of the substrate withstands the emerging shear forces. The values of the adhesion strength of ice appearing on the surface of quartz fiber heat-protective coating were obtained. It was demonstrated that the proportion of tangents in the stressed state of the tile does not exceed 10%. The proposed mathematical model for studying the characteristics of the tiled thermal protection of a spacecraft adequately describes the processes under study. Experimental studies have shown satisfactory agreement with theoretical results. The calculations performed to determine the ultimate shear stresses in the substrate showed that the strength of the substrate withstands the emerging shear forces. Consideration of climatic conditions can significantly reduce risks in the design of thermal protection of modern aircraft. The results of this study can be used in further studies, and they can also be taken into account in the construction of aircraft.

극초음속 유도탄 열보호 설계를 위한 탄화성 페놀수지 복합재의 열분해 및 화학적 삭마 해석

A chemical ablation analysis of hypersonic missile for thermal protection design was performed using SAMCEF AMARYLLIS V.17, the specific ablation module of commercial software based on a finite element code. The pyrolysis and surface recession models based on the effects of mass loss, pore gas diffusion, and endothermic reaction energy were applied for non-linearity of material and boundary. The numerical studies were carried out to confirm the tendency of the pyrolysis and chemical ablation of theoretical ablative composite for open testing (TACOT) with a chemical composition similar to charring carbon/phenolic composites. The frequency of variation in surface recession, temperature, density and gas mass flux was reviewed to determine the characteristics of multiple decomposition reactions and oxidation via pyrolysis of gas species.

Full Three-Dimensional Monte Carlo Radiative Transport for Hypersonic Entry Vehicles

A full three-dimensional photon Monte Carlo radiative transport solver, named StaRad, is developed to calculate thermal radiative heat flux and infrared signatures within equilibrium and nonequilibrium hypersonic flows. The code is developed to investigate arbitrary thermal radiative transport and is line-by-line accurate. The solver makes use of a ray plane tracing scheme. It can be used in combination with different spectral databases, such as NEQAIR, Plasma Radiation Database, or High Resolution Transmission (HITRAN)/HITEMP Radiation Database. Furthermore, it is developed to be coupled with a continuum fluid flow solver code for coupled computations. For this investigation, uncoupled computations will be presented. A validation with a reference method is presented. Because three-dimensional computation of radiative transport is time consuming, differences are investigated between piecewise one-dimensional and full three-dimensional computations. This is especially important for infrared signatures, as well as in the case of a detailed thermal protection design. This computational method can be efficiently used on a general-purpose graphical processing unit; hence, an overview of its benefits and computational speedup is given.

Test of aero-heating in hypersonic non-equilibrium flow and numerical simulation study

In the design of hypersonic vehicle thermal protection, precise prediction on the non-equilibrium aerodynamic heating in condition of hypersonic flow has been recognized as the difficult and hot topic in the current academic and engineering field. In this paper, a simple spherical column thermal model was designed. The surface catalytic properties of the model and plug-type calorimeter were changed by sputtering Au and SiO2 on the surface. In the arc tunnel, a hypersonic non-equilibrium flow aerodynamic heating test was carried out to obtain the aerodynamic heating data of the model surface under both near complete catalysis and near non-catalysis. Based on the comparative analysis of test results and numerical calculation results, the numerical prediction method of high temperature chemical non-equilibrium aerodynamic heating is validated. The results show that the catalytic properties of the surface material have a significant effect on the non-equilibrium aerodynamic heating; the catalytic effect on the spherical surface is obvious, the complete catalytic heating is higher than the non-catalytic heating, but the catalytic effect on the cylinder surface is weaker. The numerical simulation results show that the difference of the heat flux between the different wall catalytic conditions is greater than that of the ground test results; completely catalytic wall heat flux within 5% of the difference, completely non-catalytic wall heat flux difference of more than 10%.

Reaction kinetics and a physical model of the charring layer by depositing Al2O3 at ultra-high temperatures.

The thermochemical ablation of insulation material caused by slag deposition in solid rocket motors has increasingly attracted researchers' attention. Understanding the ablation mechanism and the ability to calculate reaction kinetics parameters determine the height of the thermal protection design for advanced solid rocket motors. In this work, the interaction of the Al2O3-C system is determined through static ablation experiments. Using X-ray diffraction, HSC thermodynamic software, and a thermogravimetric analyser, the carbon thermal reduction of alumina is analysed and the reaction mechanism and physical model are obtained. Isothermal experiments at 1700-1850 °C and mathematical analysis provide the kinetic parameters of the overall and step-by-step reactions. The results show that the overall reaction of the Al2O3-C system involves three steps. The overall reaction kinetics are described by the contracting area model R2 with apparent activation and frequency factors estimated as 254.5 kJ mol-1 and 5.5 × 106 min-1, respectively. The distribution reaction kinetics of steps 1 and 2 are described by the first-order chemical reaction control model (F1) and that of step 3 is described by the one-dimensional diffusion control model (D1). The corresponding activation energies are 107.9 kJ mol-1, 240.3 kJ mol-1, and 567.5 kJ mol-1, and frequency factors are 625.94 min-1, 8.3 × 105 min-1, and 1.6 × 1014 min-1, respectively.

Thermal protection study of bladder compensatory suit using a heat transfer model.

The bladder compensatory suit (BCS) is important individual protective equipment for pilots' activities in a high-flying environment. The layout and thermal diffusion ability of the bladder directly affects the thermal comfort of pilots in flight. (1) Established and verified a human-compensatory suit-environment heat transfer model; (2) Used the model to study the human thermal variation of each segment in hot conditions and clothing. To verify the two-dimensional heat transfer model, simulated data of body temperature were compared with experimental results under the same conditions (AT: 40/45°C, ordinary clothing). The model could be used to calculate the temperature variation of each body segment in three environments temperature (28°C, 35°C and 40°C) and three types of clothing (naked, ordinary clothing, BCS). The results showed that: (1) the bladder significantly affected sweating speed and skin temperature, as well as core temperature; (2) the skin temperature of the area covered by the bladder was difficult to reduce by the thermal regulation system. It was because sweat secretion was inhibited, thus, to limit evaporation. The model could be used as a reference for the thermal protection design of bladder compensatory suit. The bladder compensatory suit (BCS) is important individual protective equipment for pilots activities in a high-flying environment, and its layout directly affects the thermal comfort. Based on a two-dimensional thermal regulation system model, a body-clothing-environment heat transfer model was established. The model was used to calculate the temperature and sweat variation of each body segment in different environments and clothing.

Optimum Structural Design of Thermal Protection for Supersonic Aircraft by Using Photonic Crystal Material

With the rapid development of the supersonic aircraft technology, the aircraft Mach number continues increasing, but on the other hand, the working condition becomes progressively poor. The photonic crystals (PCs) material could reflect the energy of the thermal radiation effectively and prevent heat transferring into the substrate due to its low thermal conductivity. Consequently, the PCs material could be applied to thermal protection for the supersonic aircraft. In this paper, the aircraft state of Mach 5 is set as the target operating condition, and the PC thermal protection ability is simulated by the method of computational fluid dynamics. Based on the theory of the electromagnetics, the characteristics of the photonic band gaps for three-dimensional PCs are calculated and the effects of PCs' medium radius, refractive index, and lattice constant are fully taken into account. For the three-dimensional diamond PCs' structure, two major categories and totally five optimized design schemes are proposed, through combining the condition of supersonic aircraft aerodynamic heating. Results show that the temperature is reduced by 948.4 K when the heat passes through thermal protection layer and reduced by 930.4 K when the heat passes through PC layer. By the method of “coupled optimization strategy (COS),” the energy density which enters into substrate material would decrease by 7.99%. In conclusion, the thermal protection capacity for supersonic aircraft could be effectively improved by using the PCs.

DSMC Simulation of Hypersonic Flows over a Compression Ramp Using Subcell Scheme

Accurate prediction of the surface properties and the recirculation zone is very imperative for the design of advanced hypersonic vehicles and thermal protection. The objective of this paper is to identify the predominant factor affecting the accuracy of the calculation and to propose a more efficient way to simulate the shock wave/boundary layer interaction flow. First, the hypersonic flow over a compression ramp of 35-degree is calculated by the parallel direct simulation Monte Carlo (DSMC), and the results are compared with the benchmark to get a grid resolution. Then, the subcell scheme is implemented and the computational efficiency is highly improved. It is recommended to use a relatively coarse grid with subcell scheme to get a reliable result efficiently in engineering. It also can be found that the SOF, which is the ratio of mean collision separation (MCS) of the particles to mean free path (MFP), plays an important role in the accurate prediction of the recirculation zone, surface pressure and heat flux coefficient. Furthermore, further study show it is not a necessary condition to maintain the SOF value below unity throughout the flow field. More attention should be given to the SOF value in the cells close to the wall.