Thermal Protection System Design Research Articles

A Markov chain Monte Carlo analysis of carbon-phenolic ablation through a lumped physical model

Payloads and other rocket components use Thermal Protection Systems (TPS) to avoid the effects of high heat fluxes generated by aerodynamic heating or propellant burning. The most used type of TPS is the ablative thermal shield, employing polymeric composite ablators. Ablative and thermophysical properties of such materials must be known for TPS design. These properties are often difficult to access experimentally, and vary according to the manufacturing process, ratio between reinforcement and matrix and the fiber direction in the case of composites, and other process variables. Inverse solution methods were initially applied in computational models to solve the heat transfer problems in space vehicles. This work proposes an inverse methodology, using the Markov chain Monte Carlo (MCMC) method, to determine the ablative properties of carbon-phenolic materials manufactured by biased tape wrapping, starting from arc-jet ablation results of samples of this material. A simplified but robust lumped physical model is used to represent the direct problem. The obtained values of these properties were used in the direct problem in a more refined mesh, and the simulation was compared to the experiments. Good agreement between numerical and experimental data was obtained, allowing the use of this methodology in estimation of ablation properties.

Experimental and numerical study of the thermal response of a diesel fuel tank exposed to fire impingement

Fire exposure of diesel fuel tanks can lead to catastrophic accidents. In this study, a fire test was conducted with an 80 L cuboid fuel tank filled to 18% capacity with diesel. More importantly, an innovative three-dimensional computational fluid dynamics (CFD) model was developed aiming at predicting fuel tanks pressurization rate and temperature distributions in the inner fluid. The CFD model was verified using the experimental data and the result showed close agreement. Then, the attention was focused on the influence of different fire engulfment zones and filling levels on the thermal response of the diesel fuel tank. The results indicated that, when fire engulfed the bottom and sides of the diesel fuel tank, temperature stratification appeared in the vapor space and gradually diminished with increasing filling level; however, the liquid phase could be considered isothermal, which was dominated by the convection at the bottom. Additionally, the high filling level caused the pressurization of the fuel tank to slow down due to the absence of temperature stratification in the liquid phase. The combined experimental and numerical analysis can offer valuable guidance for improve the design of thermal protection systems and the fire risk management of fuel tank explosions.

Numerical Simulation of an Electrothermal Ice Protection System in Anti-Icing and Deicing Mode

The design of efficient thermal ice protection systems is a challenging task as these systems operate in complex environments involving several coupled physical phenomena such as phase change, boundary-layer flow, and heat transfer. Moreover, certification rules are becoming more stringent, and there is a strong incentive for the reduction of fuel consumption. In this context, numerical tools provide a powerful asset during the design phase but also to gain insight into the physical mechanisms at play. This article presents modeling and simulation strategies for thermal ice protection systems. First, the model describing the behavior of the thermal protection system is presented. Second, a model and associated numerical method is presented for unsteady ice accretion. Third, the coupling methodology between the ice accretion solver and the heat conduction solver is described. In the fourth part, different methods to simulate the boundary-layer flow are described. Finally, some relevant examples are presented, both in steady and unsteady configurations.

Dynamical behaviors of aluminum droplets impinging on horizontal and inclined surfaces

Aluminum (Al) droplets collide with the wall of the combustion chamber in a solid rocket motor at different angles of incidence and velocities, which exhibit rebounding, splashing, and other behaviors. It increases the loss of two-phase flow to degrade the performance of the engine and poses a challenge to the thermal protection system. In this study, an experimental system in which molten Al droplets impact the walls of the chamber is designed. The size of the Al droplets produced by the generator ranged from 120 μm to 2000 μm while their speeds of impact ranged from 0.85 m/s to 9 m/s. This was used to examine the process of impact of Al droplets on flat and inclined walls. The residence time and spreading diameter of the droplets were quantified to reveal their mechanism of collision. Their empirical correlations were obtained by regressing the experimental data, and the predictive error of the model was within 20% of the experimental values. Furthermore, a theoretical model was developed to analyze energy distribution during the collision between the droplets and the wall. It showed that the effect of viscous dissipation energy on droplet flow was negligibly small, but the adhesion energy of the droplet increased significantly with its initial energy. In addition, the dimensionless sliding distance was S≈8 when the tangential Weber number was greater than 90, where this was related to the high surface tension of the Al droplets. The results here provide a theoretical basis for the design of thermal protection system of solid rocket motors.

Heat transfer and pyrolysis gas flow characteristics of light-weight ablative thermal protection system in the blunt body

Thermal protection system (TPS) is the critical system to protect the structure and electronic components of hypersonic vehicles from aerodynamic heating. Since the blunt body suffers the most severe thermal environment in a flight vehicle, the design of TPS for the blunt body is crucial. To investigate the thermal performance of the light-weight ablative material in the blunt body, a multi-dimensional thermal response model is proposed considering heat transfer, pyrolysis gas flow, and gas blowing effect. The characteristics of heat diffusion, pyrolysis process, gas flow, and the effect of pyrolysis gas blowing are investigated. It shows that the developed model can accurately predict temperature and gas flow. The temperature and charred thickness at the stagnation point are the largest, decrease along the surface of the sphere nose, and keep almost the same at the cone. Due to the surface pressure distribution, a large part of the pyrolysis gas in the sphere nose flows through the ablative material and injects into the air at the sphere-cone junction. In contrast, the gas at the cone flows vertically to the cone surface. The gas flow rate and pressure decrease as time elapses. Moreover, gas blowing mainly influences the thermal response near the sphere-cone junction, leading to a minimum blowing factor of 0.965 and a max temperature reduction of 7.4% at 100s. The effect of gas blowing is enhanced with the increase of phenolic resin content. This study deepens the understanding of the thermal performance of light-weight ablative material in the blunt body and guides the ablative thermal protection design.

Kinetic Monte Carlo modeling of the interface heterogeneous catalysis effect based on machine learning

<p indent="0mm">Heterogeneous catalysis in a high-enthalpy environment has a remarkable influence on the aerodynamic heating of hypersonic aircraft, and modeling this multiscale effect is an important prerequisite for the lightweight and low-redundancy design of an aircraft thermal protection system. A mesoscale interface heterogeneous catalysis model applicable to macroscale flow field calculation was established by the Monte Carlo method. By the radial basis function neural network algorithm, a machine learning-based reaction kinetics analytical surrogate model was constructed, by which fast and accurate prediction of the catalytic recombination coefficient was realized. The results showed that the machine learning model predicted a recombination coefficient similar to the kinetic Monte Carlo method while reducing the magnitude of the calculation time to that of the traditional macroscale method. Coupled with the macroscale CFD solver, the proposed model can efficiently predict surface heat flux with a multiscale heterogeneous catalysis effect. Modeling of microscale results by the machine learning method and its application in macroscale simulation not only covers the heterogeneous catalytic kinetics but also improves computational efficiency. This modeling provides a theoretical basis for multiscale modeling of complicated high-temperature effects on hypersonic aircraft surfaces.

Application of improved k-ω-γ model to deformed hypersonic Inflatable Aerodynamic Decelerator aeroshell

The scalloping surface of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) promotes flow transition. Precise prediction of flow transition and surface aeroheating over the HIAD is imperative to the design of thermal protection system. In this paper, the improved k-ω-γ model considering the separation instability is first applied to the deformed HIAD aeroshell under various angles of attack to assess and verify its performance on such complex configuration, where multiple transition mechanisms coexist and interact. The results of the original k-ω-γ model are also provided for comparison and the prediction mechanisms of the improved model are further dissected. The results reveal that the flow separation and crossflow caused by the scalloping surface are highly sensitive to the angle of attack (α). With the α increasing, an opposite tendency between them can be found, that is, larger α causes stronger and earlier crossflow, but much later flow separation, which leads to the alternation of dominant instabilities for transition in turn. At the relatively high angle of attack, the crossflow becomes critical, at a small angle of attack, however, the separation instability is dominant. The first-mode disturbance can also accelerate the flow instability, while the influence of second-mode disturbance can be neglected due to relatively low post-shock Mach numbers. As a result, the original k-ω-γ model, not attempting to incorporate separated flow transition, predicts an opposite tendency to the experiments. While the transition onset, extent and shape simulated by the improved k-ω-γ model compare quite well with the experiments.

Transient Heat Transfer Analysis on Composite Solids for Hypersonic Flow Applications

Surface heating rates have the importance in high speed flow due to viscous or aerodynamic heating which leads to enormous rise in surface temperature. Prediction of exact heat transfer rates is essential for design of protection systems for aerodynamic bodies aimed. For the thermal protection system, the aerodynamic Surfaces are generally configured with multiple solids. Hence present study is planned to develop the conjugate heat transfer (CHT) solver for composite solids. A CHT solver is in hypersonic flow environment. Finite volume framework and explicit time stepping have been implemented for both the compressible Navier-Stokes solver and conduction solver. Loose coupling has been established between the above mentioned fluid and solid domain solvers. The following observations were made in the CHT simulations. In the case of hypersonic flows over finite thickness flat plate, considerable difference in surface heat transfer rate has been noticed along the stream wise direction at the solid-fluid interface and also the Effect of wall material on heat diffusion for hypersonic application has been analyzed using loosely coupled CHT solver. The CHT solver has also been implemented for hypersonic flow over composite cylinder. Necessity of insulating material in the region of stagnation point has been realized as the outcome of these studies.&#x0D; HIGHLIGHTS&#x0D; &#x0D; Accurate prediction of Convective heating rates is essential for the design of thermal protection system for hypersonic vehicles&#x0D; In The thermal protection system, the aerodynamic surfaces are generally configured with multiple solids. Hence the present study is planned to develop the conjugate heat transfer (CHT) solver for composite solids&#x0D; The CHT solver has been implemented for hypersonic flow over composite flat plate and cylinder. It has been observed flows over finite thickness flat plate, a considerable difference in surface heat transfer rate has been noticed along the stream wise direction at the solid-fluid interface and the necessity of insulating material in the region of stagnation point has been realized as the outcome of these studies&#x0D; &#x0D; GRAPHICAL ABSTRACT

Assessment of a 3D ablation material response model for lightweight quartz fiber reinforced phenolic composite

AbstractTo understand the ablation mechanism of materials is crucial for the design of thermal protection system (TPS) for manned reentry vehicles. To investigate the thermal behavior of lightweight quartz fiber reinforced phenolic (LQP) composite subjected to an aerodynamic hyper‐thermal environment, a 3D ablation material response model for the LQP composite is developed with surface and volume ablation in consideration. It is used to predict surface and back temperature, linear ablation velocity, and quality loss rate of the material. The 3D temperature field, pyrolysis degree, gas flow, and pressure field are analyzed as well. To evaluate the accuracy of the model, an ablation test of oxyacetylene flame is adopted on the LQP composite. In the ablation process, the pyrolysis gases propagate primarily through the surface primarily and through the sidewall afterwards. The gas pressure peaks at the early stage rather than increasing constantly. The analysis of thermal insulation mechanism of the LQP composite indicates that the radiation and heat blockage effect are the primary factors to improve the ablation resistance, accounting for 50.0% and 37.3% of the improvement, respectively. The low thermal conductivity determines the admirable thermal insulation performance. The surface temperature of material can be reduced by introducing high emissivity powder into the material, preparing coating on the surface, and adjusting the content of phenolic and silica fiber.

In-depth study on reaction kinetics of porous charred layer with high temperature alumina deposition in solid rocket motor

Current understanding of insulators abnormal ablation caused by alumina deposition in solid rocket motors (SRMs) is limited. The ability to calculate its thermochemical kinetic parameters and establish the corresponding ablation model determines the height of the thermal protection system design for advanced SRMs. In this work, the reaction mechanism of interaction between α-alumina and porous charred layer under deposition conditions was proposed firstly. Then the isothermal thermogravimetric experiments at 1750-1900 °C were carried out for the reaction system. The TG and DTG curves under various reaction conditions were treated by isoconversional method and model-fitting method, respectively, and the corresponding kinetic parameters and reaction models were calculated. Finally, based on the above research results, a deposition ablation model was established and the corresponding calculation code was compiled. Results show that the reaction of α-alumina/ porous charred layer system is formed by the superposition of four steps. The reaction model and kinetic parameters of each step reaction are as follows: steps 1 and 4 follows the diffusion control model (D3) with activation energies Ea of 109.6 kJ·mol−1, 47.9 kJ·mol−1 and frequency factors A of 47.9 min−1, 0.4 min−1, respectively, step 2 follows the chemical reaction control model (F1) with Ea of 48.5kJ·mol−1 and A of 14.8 min−1, and step 3 follows the contracting area model (R2) with Ea of 41.9kJ·mol−1 and A of 3.9 min−1. The insulator's ablation thickness calculated by the deposition ablation model established in this paper has an error of 16.01% compared to that in the experiment, indicating that it can be used to accurately calculate the ablation characteristics of insulators under the alumina deposition conditions in the future work.

Investigation of energy accommodation coefficient at gas-solid interface of a hypersonic flying vehicle

An effective thermal protection system is crucial for hypersonic flying vehicles performing multiple cycles of flight under high aerodynamic heating loads. Accurate calculation of heat transfer between a flying vehicle and its front shock waves is not possible due to complex gas-solid interactions. In this study, it was observed that the energy accommodation coefficient (EAC) at the interface between the shock wave formed by monoatomic gas and the flying vehicle exhibited a decrease-increase trend, which is attributed to that the averaged lasting time for each gas atom interacting with the solid decreases exponentially while the interacting depth increases logarithmically with the increase of the flight velocity. A one-dimensional collision model disclosed that the minimum EACs at different spring-mass systems occurred at the same Deborah number, indicating that the incident velocity of gas atoms corresponding to the minimum EAC is proportional to the natural frequency of the spring-mass system. Our research results suggest that the EAC strongly depends on the vehicle flying velocity and provide a rule for the optimal design of an efficient thermal protection system.

Catalytic characterization in plasma wind tunnels under the influence of gaseous recombination

The catalytic properties of materials used in re-usable thermal protection systems (TPSs) of re-entry vehicles are mainly characterized in plasma wind tunnels. These facilities are adequate to reproduce the thermo-chemical environment found in hypersonic flights. Catalytic recombination coefficients (W) determined empirically for the TPS design process are used to validate catalytic models. Recent experiments in plasma wind tunnels present some contradictions with current catalytic models used in computational fluid dynamics simulations. This work explores the coupling between homogeneous and heterogeneous reactions and its influence on the heat flux measurements, in order to explain such contradictions. For that, three copper calorimeters (at 350 K) are inserted in probes of different sizes, and they are tested in the Plasmatron facility at the von Karman Institute under different chemical regimes, which are modified by the probe diameter. A dimensional analysis of the acquired data based on previous non-equilibrium boundary formulations shows that an empirical recombination coefficient can be falsified by the presence of homogeneous reactions. Then, in the framework for the post-flight analysis of the European Space Agency's intermediate experimental vehicle, the local heat transfer simulations methodology is applied under peak heating conditions. A new probe is designed, manufactured, and tested in the Plasmatron to measure the recombination coefficient on C/SiC under the closest flow conditions with respect to the flight.

Thermal insulation and anti-vibration properties of MoSi2-based coating on mullite fiber insulation tiles

Thermal protection materials with excellent thermal insulation properties and high reliability are crucial for aerospace vehicle. Mullite fiber insulation tiles coated with MoSi2-borosilicate glass ([email protected]2) were prepared by a simple slurry method. Results shown that the surface temperature rapidly reached 1043.1 °C under the heat flux of 450 kW/m2, while the cold-surface remained at room temperature. The adjusting effects of MoSi2-based coating on the thermal response and transfer properties of MFIT were investigated systematically. Compared with MFIT, the surface and internal temperatures of [email protected]2 were obviously suppressed, due to the existence of MoSi2-based coating with high emissivity, which effectively enhanced the thermal radiation of the surfaces. Finally, the structural reliability under coupled environment of heating (q0 = 450 kW/m2) and random vibration with high magnitude (20–2000 Hz, Grms = 20 g) was also investigated. Destructive cracking and peeling of the coating were not observed, due to the excellent bonding strength between the coating and the substrate. The results provided an important basis for the potential application of insulation tiles with mullite fibers in aerospace vehicles and the design of lightweight thermal protection systems.

Integrated Thermal Protection System Design for Hypersonic Vehicle Based on New Thermal–Mechanical Method

With the development of hypersonic vehicles, considerable attention has been paid to thermal protection systems (TPSs). Among the different types of TPS, integrated TPS (ITPS) has attracted extensive interest because of its light weight, cost effectiveness, bearing capacity, and easy maintenance. However, ITPS, with its complex structure and reusable requirement, is facing challenges and it is worth devoting much effort to this. In an effort to overcome these challenges, we develop a novel approximate analytical method based on separation of variables and orthogonal expansion technique, which is presented for the prediction of heat transfer. The approximate analytical method has the ability to consider the effects of temperature-dependent thermal material properties, convection and radiation. Moreover, a C0 higher-order layer-wise finite-element model combined with homogenization techniques and a simplified three-dimensional (3D) finite-element model based on the periodic structure characteristic are proposed to estimate the thermal–mechanical response and stability of ITPS effectively. In addition, an optimization procedure based on the proposed methods for ITPS is developed. Implementation of the optimization is demonstrated by applying it to design the ITPS of space shuttles and hypersonic vehicles. By comparing with conventional thermal protection systems (CTPSs), ITPS exhibits an outstanding advantage in weight and an acceptable disadvantage in size. The objective of this paper is to establish an accurate and efficient design method aiming at determination of the optimal performance of ITPS, and puts forward the foundation for the development of a hypersonic vehicle.

Thermo-mechanical characterization of carbon-based ceramic foams for high temperature space application

The successful development of aerospace reusable launch vehicles (RLV) require to realize effective thermal protection systems (TPS) for preserving spacecraft integrity from the severe thermal loads during re-entry phase. To such an aim, due to the need of reducing payload transportation costs, applied research is driven towards lightweight materials with advanced thermo-mechanical properties. Space TPS are often based on sandwich structures, where the core material has the main function of thermal insulation. Ceramic porous materials, as carbon (C) and silicon carbide (SiC) foams, represent ideal candidates for application as structural TPS component, thanks to both low density and significant thermal stability at very high temperatures. The paper presents a joint experimental study of carbon-based ceramic foams proposed as sandwich’s core for TPS design. A full thermal characterization of commercial C- and SiC-foam materials is reported, including measurements of thermo-mechanical combined stress, temperature-induced outgassing behavior and heat transfer properties. These latter, in particular, are studied by means of a robust numerical technique, known as the inverse method, which allows to evaluate materials thermal conductivity and heat capacity over a wide range of temperatures, thus establishing the required material behavior for potential use in spacecraft TPS.

Advanced sandwich structures for thermal protection systems in hypersonic vehicles: A review

A heat shield called the thermal protection system (TPS) is an important structure in hypersonic vehicles as it prevents hot air from entering vehicles and potential impacts from space debris. With the increase in demand for low-cost reusable launch vehicles as well as for searching and exploration of new planets in both unmanned and manned missions, the need for developing an effective TPS has increased across many countries. The structural design of TPSs has become more prominent in the early stage of hypersonic vehicle development. Sandwich structures that have the advantages of low density and high performance are integrated into the structural design of an effective TPS. This paper provides a comprehensive review of recent research efforts on sandwich structures for TPSs. The topics discussed in this paper include aspects of structural and material design, mechanical and thermomechanical performances, and manufacturing methods. In particular, we review and discuss the structural design as well as the material design of sandwich structures for different TPS types with various configurations, including corrugated cores, lattice cores, multilayer cores, foams, honeycomb cores, bio-inspired cores. The materials used for the sandwich structures, such as various types of laminated composite, ceramic matrix composite, and metals, are included. We also discuss the performance of the TPS sandwich structures in terms of temperature gradients, deformation limits, and mechanical strengths and provide a discussion on the manufacturing methods of TPS sandwich structures for hypersonic vehicles. Finally, further research directions and challenges of sandwich structures for TPSs are presented.

Conjugate heat transfer study of hypersonic flow past a cylindrical leading edge composed of functionally graded materials

Aero-thermodynamic analysis of a cylindrical leading edge placed in a hypersonic stream is carried out using an in-house developed conjugate heat transfer (CHT) solver. Isotropic and functionally graded materials (FGM) are tested as heat shields to understand the effects of the material property on the flow structure and aerodynamic heating associated with the mutual coupling of fluid flow and heat transfer. A simplified partitioned approach is employed to couple the independently developed fluid flow and heat transfer solvers to perform conjugate heat transfer studies. This framework employs a cell-centred finite volume formulation with an edge-based algorithm. Both strong and loose coupling algorithms are implemented for the data transfer across the fluid–solid interface. A test case of hypersonic flow over a cylindrical leading edge composed of an isotropic material is considered to validate the accuracy and correctness of numerical formulation adopted in the in-house solver. The significance of solid domain materials on the conjugate heat transfer has been studied by considering both isotropic material and FGM. The loosely coupled CHT solver required 10 times less simulation time when compared with the strongly coupled CHT solver. The interface heat flux evolution over time showed a decreasing trend, whereas an increasing trend was for the interface temperature. The current study strongly recommends CHT analysis for the design of thermal protection system of space vehicles. The thermal performance of FGMs composed of various volume fractions of Zirconia and Titanium alloy (Ti6Al4V) is assessed. The temperature distributions obtained from the CHT analysis shows that FGM with a power index of unity is a good material choice for thermal protection systems.

Influence of ambient pressure on thermal response of polymer matrix composites

AbstractAmbient pressure is crucial to the thermal performance of polymer matrix composites, which are usually used as the thermal protection system (TPS) materials of hypersonic vehicles. Aiming at studying the influence of ambient pressure on thermal response of polymer matrix composites, a novel pyrolysis layer model with pressure is presented, and its mathematical model is deduced and programmed in MATLAB. Comparison between numerical results and experimental results of theoretical ablative composite for open testing is given to verify the accuracy of the presented model. After that, thermal responses of phenolic impregnated carbon ablator under ambient pressures range from 1 to 500 kPa are simulated. It is found that accumulation and backstreaming of generated pyrolysis gasses begin to appear with the increase of ambient pressure. And there is an obvious linear correlation between the reciprocal of the material internal pressure and the ambient pressure. In addition, bondline temperature of polymer matrix composites declines as the ambient pressure decreases. Effect of ambient pressure on thermal response of polymer matrix composites cannot be neglected, which should be taken into account in the TPS design. This study contributes to the accurate prediction of thermal response of polymer matrix composites.

InSight Aerothermal Environment Assessment

The Mars Interior Exploration using Seismic Investigations, Geodesy and Heat Transport (InSight) spacecraft, which successfully touched down on the planet surface on November 26, 2018, was proposed as a near build-to-print copy of the Mars Phoenix vehicle to reduce the overall cost and risk of the mission. Because the lander payload and the atmospheric entry trajectory were similar enough to those of the Phoenix mission, it was expected that the Phoenix thermal protection material thickness would be sufficient to withstand the entry heat load. However, allowances were made for increasing the heatshield thickness because the planned spacecraft arrival date coincided with the Mars dust storm season. The aftbody thermal protection system components were not expected to change. In a first for a U.S. Mars mission, the aerothermal environments for InSight included estimates of radiative heat flux to the aftbody from the wake. The combined convective and radiative heat fluxes were used to determine if the as-flown Phoenix thermal protection system design would be sufficient for InSight. Although the radiative heat fluxes on the aftbody were predicted to be comparable to, or even higher than the local convective heat fluxes, all analyses of the aftbody thermal protection system showed that the design would still be adequate. Aerothermal environments were computed for the vehicle from postflight reconstruction of the atmosphere and trajectory and compared with the design environments. These comparisons showed that the predicted as-flown conditions were less severe than the design conditions.

Thermal Environment of Launch Pads During Rocket Launching

This paper investigates the thermal environment of launch pads during rocket takeoff. The rocket plume model is set up using the three-dimensional Navier–Stokes equations and realizable turbulence model. The aluminium oxide particles in the plume are considered by the Eulerian dispersed phase model. Comparing with the experimental data, the accuracy of the numerical model can be verified. On this basis, in total 16 numerical cases are performed to analyze the influence of the rocket flight altitude and lateral drift on the temperature of the launch pad. The results show that the launch pad has a more severe thermal environment at the flight altitudes from 3 to 20 m. Because of the rocket drift, a high-temperature gas flow layer is formed on the launch pad, which results in a dramatic increase in the temperature of the cross beam. At this location, the maximum temperature can reach up to 3900 K, 30% higher than the value without rocket drift. The method in this paper can provide an effective way to evaluate the thermal environment of the launch pad during rocket launching and have a great value on the thermal protection system design.