Heat Shield Design Research Articles

Optimal design of selected features of exhaust system shields using different optimization methods and artificial neural networks

The article presents problems related to methods of optimal design of heat shields used in exhaust systems of internal combustion engines. The optimization method proposed in this paper goes well beyond the scope of the standard design process. The paper uses a variety of local and global optimization algorithms, those both built into numerical simulation systems and in-house and external algorithms. An optimization criterion was defined and numerically implemented, together with constraints derived from the real requirements for this type of shielding. A simplified numerical model of finite element method providing the required accuracy adapted to the optimization task was developed. In addition, the work also presents a method for creating finite element surrogate models using artificial neural networks. The process of selecting the network topology and its learning allowed the development of a metamodel characterized by very good quality, for which, despite the relatively large number of design variables, the response errors are completely acceptable from a practical point of view. Numerical results were compared and developed for the used methods and algorithms.

The effects on heat and oxygen transport of different heat shield and sidewall insulation designs during continuous Czochralski silicon crystal growth

Numerical work has been done to optimize the hot zone for the continuous Czochralski growth of 8-inch diameter silicon crystals using a double-crucible system. Thicker insulation layers on a high-heel-shaped heat shield and the sidewall insulation led to a reduction in the heater power of more than 40%, a decrease in the oxygen content, and a significant flattening of the growth interface. This design also allowed a stronger argon gas flow near the outer free melt surface and the quartz crucible wall, which prevented the generation of an argon gas vortex in this area. Stronger flow of argon gas would assist in moving more SiO gas and dust out of the chamber. However, the reduction in the outer melt temperature compared to the original design could slow down the melting of the silicon granules.

Performance of cork-based thermal protection material P50 exposed to air plasma

Cork P50 thermal protection material was characterized under the Earth atmospheric entry conditions in the framework of the in-flight experiment QARMAN. A total of 25 P50 samples were exposed to air plasma ranging the chamber static pressure values of 1500, 4100, 6180, and 20,000 Pa and heat fluxes between 280 and 3250 kW/m^2. The sample radius was also varied between 11 and 25 mm radius to investigate the effects of the stagnation line velocity gradient. The experimental results on heat flux, pressure, surface and in-depth temperatures, surface emissivity, swelling and recession rates, and computed boundary layer profiles as well as the mass blowing rates are presented. A material characterization data suite is also provided including thermogravimetric analysis, specific heat, and thermal conductivity. Overall, a wide range of experimental data of Cork P50 material in Earth atmospheric entry conditions is made available to modelers and engineers for improved material response models and heat shield design consolidation.

Finding Ablation Rates via Temperature Measurements from the Mars Science Laboratory

The determination of the ablation rate for thermal protection systems on atmospheric entry is a difficult task, but can have significant implications in terms of safety and heat shield design. Ablation sensors of various kinds can be used to measure ablation directly, but in the event of a failure of these sensors, the use of thermocouples imbedded in the heat shield for determining the ablation rate is somewhat more difficult. A numerical model was developed as part of this research to determine ablation rate as a function of temperature using measurements from the Mars Science Laboratory Entry, Descent, and Landing Instrumentation Mission, which took place in 2012. The numerical model accounts for ablation as well as temperature-dependent thermal conductivity and volumetric heat capacity. A temperature-dependent ablation rate is calculated simultaneously with the heat flux for each time step of data acquisition. Regularization is provided for the heat flux calculation through the use of cubic splines. A significant reduction in the residual values between the mathematical model and the measured temperatures was obtained when accounting for ablation and refined temperature-dependent thermal properties. In some cases, the standard deviation of the residuals was reduced by 70% or more, indicating a greatly improved conformance of the mathematical model to the measured temperatures when accounting for ablation.

Modeling Heat-Shield Erosion due to Dust Particle Impacts for Martian Entries

Heat-shield design for spacecraft entering the atmosphere of Mars may be affected by the presence of atmospheric dust. Particle impacts with sufficient kinetic energy can cause spallation damage to the heat shield that must be estimated. The dust environment in terms of particle size distribution and number density can be inferred from ground-based or atmospheric observations at Mars. Using a Lagrangian approach, the particle trajectories through the shock layer can be computed using a set of coupled ordinary differential equations. The dust particles are small enough that noncontinuum effects must be accounted for when computing the drag coefficient and heat transfer to the particle surface. Surface damage correlations for impact crater diameter and penetration depth are presented for fused silica, AVCOAT, shuttle tiles, cork, and Norcoat® Liège. The cork and Norcoat Liège correlations are new and were developed in this study. The modeling equations presented in this paper are applied to compute the heat-shield erosion due to dust particle impacts on the ExoMars Schiaparelli entry capsule during dust storm conditions.

Numerical Methodology for the Conceptual Design of Conformal Ablative Heat Shields

Conformal ablators are low-density composite materials comprising a flexible fibrous substrate and polymer matrix. Recent advancements have improved the efficiency of conformal ablator fabrication through vacuum infusion processing, which injects resin into a fiber substrate enclosed in a matched mold under vacuum. A conceptual design methodology was developed combining a Darcy’s law simulation of the mold filling process with optimal tile layout and material property estimation. The framework allows rapid evaluation of changes to constituent materials, process parameters, and heat-shield geometry, and their impact on manufacturing and material properties. Results for a 4.5 m, 70 deg sphere-cone aeroshell demonstrate the advantages of the integrated approach. The mold filling simulation is then compared with experimental results.

Downrange manoeuvre and oscillation suppression of a self-regulating centrifugally deployed flexible heat shield using a controlled reaction wheel

A recent study has introduced a flexible deployable heat shield that passively deploys and stiffens due to centrifugal forces generated from a self-regulated autorotation. This paper demonstrates that the heat shield is similar to a PI controlled second order nonlinear system, which explains why the deployment is accompanied by a limit cycle structural oscillation that persists throughout a simulated re-entry. The heat shield design offers a unique capability to actively adjust the deployment using conventional attitude control devices. This operation is explored by simulating the re-entry of a CubeSat-sized vehicle equipped with an off-the-shelf reaction wheel controlled by a switching phase shift controller and gain-scheduled controllers. The effects of the control parameters are investigated, and successful oscillation suppression as well as an open-loop downrange manoeuvre of over 300 km is predicted for re-entry from low earth orbit.

Numerical study on the aerothermodynamics of different heatshield configurations for Mars entry capsules

Aeroheating predictions play an important role in the heatshield design of Mars entry capsule. This paper numerically investigates the effects of sphere-cone angle on the aerothermodynamic performances of heatshield configurations. Three-dimensional Navier-Stokes equations with chemical non-equilibrium models are employed to simulate the flowfield around the capsule. The laminar and turbulent heating rates of different heatshield configurations are compared and analyzed in detail. A novel correlation for turbulent heating augmentation in terms of laminar momentum thickness Reynolds number is developed to provide a guidance for engineering design and application. The proposed correlation can be more accurate and applicable due to its consideration of the sphere-cone angle effects. Finally, the laminar and turbulent maximum heat flux and total heat load along the flight trajectory are investigated for all the configurations. The numerical study is expected to illustrate the aeroheating characteristics of different heatshield configurations and provide an insight into the rational configuration design for future Mars entry capsules.

Calibration and Uncertainty Quantification of VISTA Ablator Material Database Using Bayesian Inference

The current design of NASA’s Multi-Purpose Crew vehicle uses the latest iteration of AVCOAT, an ablating thermal protection material. However, restrictions placed on the export of experimental data concerning its performance make collaborative efforts aimed at improving existing heat-shield design tools difficult to establish. The material model dubbed VISTA provides an alternative open-source platform upon which the physics of ablation can be thoroughly investigated. In this paper, calibration of a material model through Bayesian inference is demonstrated with VISTA and open-source Apollo-era AVCOAT flight data. A sensitivity analysis is first carried out using Pearson correlation coefficients and method of Sobol where the results of both approaches are compared. The calibration methodology used in this paper is then demonstrated first with the use of manufactured data. Following, the parameters of VISTA are calibrated using material temperature data recorded during the Apollo 4 test flight and uncertainties due to parametric, modeling, and data inaccuracy sources are simultaneously quantified. Uncertainty quantification of model output in this work is done by forward propagating quantified uncertainties onto model output where a large reduction in total uncertainty is observed.

Sensitivity analysis of geometric parameters upon the aerothermodynamic performances of Mars entry vehicle

Aerothermodynamic environment predictions play an important role in the heatshield design of Mars entry vehicle. This article investigates and presents the influences of geometric parameters of the heatshield on its aerothermodynamic performances. The three-dimensional coupled implicit compressible Reynolds Averaged Navier–Stokes (RANS) equations and perfect gas model with the specified effective specific heat ratio have been applied to numerically simulate the flow fields around the vehicle. Menter’s shear stress transport (SST) turbulence model with compressible correction is implemented to take account of the turbulent effect. The laminar and turbulent heating rates are demonstrated and analyzed in detail. Furthermore, a non-intrusive polynomial chaos (NIPC) method with Latin hypercube sampling (LHS) is utilized to establish the functional relationship between the aerothermodynamics and geometric parameters. In addition, Sobol indices as global sensitivity metrics have been introduced to investigate the relative contribution of each geometric parameter. The results show that for the maximum heat flux, the value of the cone angle (αc) with a high index is the top contributor to the both laminar and turbulent flow state, thus the geometric parameter αc should be considered firstly in the material design process of thermal protection system. Moreover, in the most region of MSL heatshield, cone angle (αc) also became the major influence factor. However, in a relatively small region, aerothermodynamics exhibits a great sensitivity to the change of nose radius (Rn). In all regions of heatshield, the parameter of shoulder radius (Rs) is always at a low level of Sobol index.

Optimum design of ablative thermal protection systems for atmospheric entry vehicles

The Thermal Protection System (TPS) provides spacecrafts entering the atmosphere with the thermal insulation from the aerothermodynamic heating. The design of such a subsystem is very critical, considering that its damage can lead to a catastrophic failure of the whole entry system, in particular if ablative materials are considered. In order to design an ablative TPS, in fact, a reliable numerical procedure, able to compute surface recession rate, pyrolysis and internal temperature histories under severe heating conditions, is necessary. Indeed, the TPS needs to be sized to effectively shield the spacecraft from the high heat fluxes acting during the atmospheric entry phase. At the same time, its weight has to be the minimum value able to guarantee a suitable protection.This article aims to describe an optimization procedure for the design of ablative heat shields. In particular, in the present work, the numerical method is applied to the ablative TPS of the hypersonic reentry capsule Stardust.

A generic local thermal equilibrium model for porous reactive materials submitted to high temperatures

Many engineering applications involve reacting porous materials submitted to high temperatures. This work presents a detailed but pragmatic heat and mass transfer model for porous materials containing several solid phases and a single gas phase. The detailed chemical interactions occurring between the solid phases and the gas phase are modeled at the pore scale assuming local thermal equilibrium. Homogenized models are obtained for solid pyrolysis, pyrolysis species injection in the gas phase, heterogeneous reactions between the solid phases and the gas phase, and homogeneous reactions in the gas phase. The chemistry models are integrated in a macroscopic model making use of volume-averaged governing equations for the conservation of solid mass, gas mass, species (finite-rate chemistry) or elements (equilibrium chemistry), momentum, and energy. The model has been implemented in the Porous Material Analysis Toolbox based on OpenFoam (PATO), distributed Open Source by NASA. Applications to two high-temperature engineering problems are presented. The first application concerns the design of heat-shields of space vehicles. The second one aims at improving the understanding of biomass pyrolysis for the production of biohydrocarbons.

Control of heat transfer in continuous-feeding Czochralski-silicon crystal growth with a water-cooled jacket

The continuous-feeding Czochralski method is an effective method to reduce the cost of single crystal silicon. By promoting the crystal growth rate, the cost can be reduced further. However, more latent heat will be released at the melt-crystal interface under a high crystal growth rate. In this study, a water-cooled jacket was applied to enhance the heat transfer at the melt-crystal interface. Quasi-steady-state numerical calculation was employed to investigate the impact of the water-cooled jacket on the heat transfer at the melt-crystal interface. Latent heat released during the crystal growth process at the melt-crystal interface and absorbed during feedstock melting at the feeding zone was modeled in the simulations. The results show that, by using the water-cooled jacket, heat transfer in the growing crystal is enhanced significantly. Melt-crystal interface deflection and thermal stress increase simultaneously due to the increase of radial temperature at the melt-crystal interface. With a modified heat shield design, heat transfer at the melt-crystal interface is well controlled. The crystal growth rate can be increased by 20%.

Coupled versus uncoupled optimal solutions for thermal and dynamic problems in spacecraft atmospheric flight

Purpose 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. Design/methodology/approach 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. Findings 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. Originality/value 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.

The First Mission of Iranian Spacecraft: Heat Shield Design

In the present paper, the Iranian spacecraft is considered, and is studied from an aero thermo dynamic viewpoint to design the proper thermal protection system, for its first mission: the suborbital flight. The design procedure is based on the decoupling of the aerodynamic heating phenomena and the solid phase heating process. Due to the sub-orbital mission of this spacecraft as a first flight, an elastomeric material is used as an unconventional heat shield to protect the blunt body reentry vehicle. The full Navier-Stokes equations are used for heat-flux estimation during flight, and the one-dimensional equations for an ablative materials are used to design the thermal protection system, according to the structural criteria.

Minimal Failure Probability for Ceramic Design Via Shape Control

We consider the probability of failure for components made of brittle materials under one time application of a load, as introduced by Weibull and Batdorf-Crosse. These models have been applied to the design of ceramic heat shields of space shuttles and to ceramic components of the combustion chamber in gas turbines, for example. In this paper, we introduce the probability of failure as an objective functional in shape optimization. We study the convexity and the lower semi-continuity properties of such objective functionals and prove the existence of optimal shapes in the class of shapes with a uniform cone property. We also shortly comment on shape derivatives and optimality conditions.

Ablative Heat Shield Design for Reentry Vehicle Using Numerical Analysis

At this paper, the thermal behavior of ablative materials as heat shields for reentry vehicles is investigated numerically. A one-dimensional finite difference solver is developed to simulate governing mass and energy equations. Four ablative materials; AVCO 5026-HCG, Carbon-Phenolic, Nylon-Phenolic, and Silica-Phenolic; are considered as a heat shield material for a reentry capsule with the diameter 2.8 meter. A heat flux profile from a simulated trajectory of a reentry capsule is used for investigation the performance and essential thickness of these four ablative materials. The only restriction for this simulation is the ultimate temperature of the backup structure which is beneath the ablative heat shield. At all simulations, the final thickness is defined by reaching the interface temperature, the temperature of the border between the ablative heat shield and solid shell, to 80±0.5 degrees Celsius. In addition, the sensitivity analysis is carried out to investigate the effect of properties variations on Carbon-Phenolic’s thickness for this specific heat flux profile.

Arcjet Testing in Shear Environment for Mars Science Laboratory Thermal Protection System

Thermal protection system materials for the Mars Science Laboratory mission were tested in the NASA Ames 60 MW arcjet in a shear environment. Shear tests were performed on candidate ablative heat shield materials in wedge and swept cylinder test fixtures. In portions of the expected flight environment, the proposed main heat shield material, known as SLA-561V, recessed orders of magnitude faster than predicted. An alternate main heat shield material, known as phenolic impregnated carbon ablator, behaved reasonably well in all regions of the flight envelope investigated here. However, the measured recession rate of the phenolic impregnated carbon ablator was on average 50% greater than the predicted recession by the fully implicit ablation and thermal response code, and in some cases the measured recession was as much as 150% greater (uncertainties included). Phenolic impregnated carbon ablator’s higher recession rate than predicted (in shear) resulted in adding thickness margin to the flight heat shield design. Other tests of the phenolic impregnated carbon ablator heat shield system, including tests of various gap fillers, damage scenarios, and repair scenarios, demonstrated the robustness of this heat shield system for use in the predicted Mars Science Laboratory flight environment.

Vehicle Hot Surface Ignition and Mitigation Measures of R-1234yf Refrigerant for MAC Systems

The European Commission (EC) as well as the United States Environmental Protection Agency (EPA) published legislations to regulate or encourage the use of low Global Warming Potential (GWP) refrigerants applied to Mobile Air Conditioning (MAC) systems. Europe mandates a GWP less than 150 of MAC refrigerants for new vehicle types. The thermodynamic refrigerant properties of R-1234yf are slightly different from the properties of R-134a, currently used in MAC systems. Although the basic material data show that R-1234yf is flammable, ignition tests performed for an automotive engine under-hood environment reveal design and packaging influences of its ignition behavior. After extensive collaborative research in 2009, the Society of Automotive Engineers Cooperative Research Team (SAE CRP1234) concluded that R-1234yf is suitable for use in automotive applications. Further ignition risk assessment regarding R-1234yf usage in MAC systems was done by SAE CRP1234-4 in 2013. They concluded that are still very small compared to the risks of a vehicle fire from all causes and well below risks that are commonly viewed as acceptable by the general public. This paper will compare under-hood hot surface ignition behavior of five different vehicles under specific test conditions. Out of these five, two vehicles were further investigated and the engine compartment designs were assessed. A broad range of flammability testing was completed to study release of the R-1234yf and Polyalkylene Glycol (PAG) oil mixture. Both vehicles were equipped with a transverse gas engine, front take down exhaust routing, turbo charger, and single evaporator MAC system. This engine architecture was chosen due to the short distance from a potential condenser leak and the hot exhaust system components in case of a frontal crash. Although the engine architecture was the same for both vehicles, the engine design and packaging were different and were described in this paper. Parameters that were considered include heat shield design, engine packaging, exhaust system surface temperatures during refrigerant release, and the impact on the time to ignition of the R-1234yf/PAG mixture. Mitigation measures were also identified and their potential integration into a vehicle were discussed. As engine coolant will be released in many real-world frontal crash scenarios, the impact of vaporized glycol water mixture on the ignition behavior was assessed. This paper further describes the MAC systems, the test facility used, and the testing procedure applied. Language: en

Monte Carlo Approach to FIAT Uncertainties with Applications for Mars Science Laboratory

A new scriptedMonte Carlo analysis software tool has been developed for use with the fully implicit ablation and thermal response program thermal-protection-system-sizing code. This paper describes the new software tool and gives examples of its increased utility and useful insights onmany aeroshell heatshield design considerations, such as maximum bond line temperature, recession, isotherm depth, surface temperature, and char depth. Finally, results are presented that illustrate the effect of uncertainties in key parameters that were used in the Mars Science Laboratory aeroshell thermal protection system design. The parameters having the greatest influence on bond line temperature dispersion for theMars Science Laboratory backshell are virginmaterial thermal density (34%), initial material temperature (20%), virgin material conductivity (17%), and char density (13%).