Ablation and Thermal Response Program for Spacecraft Heatshield Analysis
01 May 1999-Journal of Spacecraft and Rockets (American Institute of Aeronautics and Astronautics (AIAA))-Vol. 36, Iss: 3, pp 475-483
TL;DR: An implicit ablation and thermal response program for simulation of one-dimensional transient thermal energy transport in a multilayer stack of isotropic materials and structure which can ablate from a front surface and decompose in-depth is presented in this article.
Abstract: An implicit ablation and thermal response program is presented for simulation of one-dimensional transient thermal energy transport in a multilayer stack of isotropic materials and structure which can ablate from a front surface and decompose in-depth. The governing equations and numerical procedures for solution are summarized. Solutions are compared with those of an existing code, the Aerotherm Charring Material Thermal Response and Ablation Program, and also with arcjet data Numerical experiments show that the new code is numerically more stable and solves a much wider range of problems compared with the older code. To demonstrate its capability, applications for thermal analysis and sizing of aeroshell heatshields for planetary missions, such as Stardust, Mars Microprobe (Deep Space n), Saturn Entry Probe, and Mars 2001, using advanced light-weight ceramic ablators developed at NASA Ames Research Center, are presented and discussed.
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TL;DR: In this article, a review of the state-of-the-art efforts on polymeric ablative materials is presented, starting from the state of the art solutions currently used as TPS, up to covering the most recent efforts for nanostructuring their formulations.
268 citations
TL;DR: In this paper, the authors provide a comprehensive survey on the technology development of different classes of TPS from mid- twentieth century to the present time, including passive, semi-passive and active systems.
158 citations
TL;DR: In this paper, an ablation and thermal response model was developed for newly manufactured material, including emissivity, heat capacity, thermal conductivity, elemental composition, and thermal decomposition rates.
Abstract: Phenolic Impregnated Carbon Ablator was the heatshield material for the Stardust probe and is also a candidate heatshield material for the Orion Crew Module. As part of the heatshield qualification for Orion, physical and thermal properties were measured for newly manufactured material, included emissivity, heat capacity, thermal conductivity, elemental composition, and thermal decomposition rates. Based on these properties, an ablation and thermal-response model was developed for temperatures up to 3500 K and pressures up to 100 kPa. The model includes transversely isotropic and pressure-dependent thermal conductivity. In this work, model validation is accomplished by comparison of predictions with data from many arcjet tests conducted over a range of stagnation heat flux and pressure from 107 W/cm 2 at 2.3 kPa to 1100 W/cm 2 at 84 kPa. Over the entire range of test
153 citations
TL;DR: In this paper, the authors present a comprehensive literature review on three active cooling methods, i.e., regenerative cooling, film cooling, and transpiration cooling, including the fluids flow, heat transfer, and thermal cracking characteristics of different hydrocarbon fuels.
148 citations
TL;DR: The Porous-material Analysis Toolbox as mentioned in this paper is a fully portable OpenFOAM library that is implemented to test innovative multiscale physics-based models for reacting porous materials that undergo recession.
Abstract: The Porous-material Analysis Toolbox based on OpenFOAM is a fully portable OpenFOAM library. It is implemented to test innovative multiscale physics-based models for reacting porous materials that undergo recession. Current developments are focused on ablative materials. The ablative material response module implemented in the Porous-material Analysis Toolbox relies on an original high-fidelity ablation model. The governing equations are volume-averaged forms of the conservation equations for gas mass, gas species, solid mass, gas momentum, and total energy. It may also simply be used as a state-of-the-art ablation model when the right model options are chosen. As applications, three physical analyses are presented: 1) volume-averaged study of the oxidation of a carbon-fiber preform under dry air, 2) three-dimensional analysis of the pyrolysis gas flow in a porous ablative material sample facing an arcjet, and 3) comparison of a state-of-the-art and a high-fidelity model for the thermal and chemical respo...
129 citations
References
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TL;DR: In this article, the authors developed a high-level methodology for predicting entry of the Stardust sample return capsule with coupled radiation and ablation using a Navier-Stokes solver.
Abstract: The development of a new high- delity methodology for predicting entry ows with coupled radiation and ablation is described. The prediction methodology consists of an axisymmetric, nonequilibrium, Navier–Stokes ow solver loosely coupled to a radiation prediction code and a material thermal response code. The methodology is used to simulate the 12.6-km/s Earth atmospheric entry of the Stardust sample return capsule using ablating and nonablating boundary conditions. These ow simulations are used to size and design the Stardust forebody and afterbody heatshields and develop arcjet test conditions and models. The ow simulations indicate that the afterbody heating and pressure pro les in time are signi cantly different than the forebody heating and pressure pro les. This result is explained in terms of the pertinent aerothermodynamicsof the ow eld and the vehicle’s geometry.When applied to the afterbody thermal protection system, these results show that the traditional afterbody heatshield design approach is nonconservative for the Stardust sample return capsule shape and entry conditions.
254 citations
17 Jun 1996
TL;DR: In this article, the authors present the development of the light weight Phenolic Impregnated Carbon Ablators (PICA) and its thermal performance in a simulated heating environment for planetary entry probes.
Abstract: This paper presents the development of the light weight Phenolic Impregnated Carbon Ablators (PICA) and its thermal performance in a simulated heating environment for planetary entry probes. PICA material was developed as a member of the Light Weight Ceramic Ablators (LCAs) family, and since then, the manufacturing process of this material was significantly unproved. The density of PICA material ranges from 0.224 to 0.321 g/cc having uniform resin distribution within the fibrous substrate. Surface densification was also developed to improve the ablation characteristics of PICA against extremely high stagnation pressures. The thermal performance of PICA was evaluated in the Ames arc jet facility at cold wall heat fluxes from 425 to 3360 W/cm and surface pressures of 0.1 to 0.43 attn. Heat loads used in these tests varied from 6,245 to 33,600 J/cm and are representative of the entry conditions of several proposed Discovery missions. Surface and in-depth temperatures were measured by using optical pyrometers and thermocouples. Surface recession was also measured by using a template and a height gage. The ablation characteristics and efficiency of the PICA is quantified by using the effective heat of ablation, and the thermal penetration response is evaluated by the thermal soak data. In addition, comparison of the thermal performance of standard and surface densified PICA is also discussed.
221 citations
01 Jun 1968
TL;DR: In this article, a finite difference equation analysis of in-depth response of materials exposed to high temperature environment is presented. But the analysis is limited to finite difference equations. But it is not restricted to a single material.
Abstract: Computer program for finite difference equation analysis of in-depth response of materials exposed to high temperature environment
162 citations
TL;DR: The Mars Pathe nder probe contained instrumentation that measured heatshield temperatures during entry and an analysis of the entry environment and material response are presented in this paper, where Navier and Stokes forebody heating calculations show a peak unblown radiative-equilibrium heat e ux of 118W/cm 2 at the stagnation point and120 W /cm 2 on the shoulderforturbulente ow.
Abstract: The Mars Pathe nder probe contained instrumentation that measured heatshield temperatures during entry. A description of the experiment, the data, and an analysis of the entry environment and material response are presented. Navier ‐Stokes forebody heating calculations show a peak unblown radiative-equilibrium heat e ux of 118W/cm 2 at thestagnation point and120 W /cm 2 on theshoulderforturbulente ow. Theheatload is3.8 kJ /cm 2 on thenose,decreases along thefrustum,then increasesto 2.7 ‐3.1kJ/cm 2 on theshoulder depending on the onset time forturbulence. One-dimensional charringmaterialresponseiscalculated using threedifferentmodels.Stagnationpoint temperature data are consistent with about 85% of fully catalytic laminar heating. Shoulder temperature data are inconclusive, but are consistent with fully catalytic laminar heating or with 85% of fully catalytic heating with early onset of turbulence. Aft temperature data indicate a peak heat e ux and heat load of about 1.3 W /cm2 and 70 J/cm 2 , respectively. The aft heating proe le is about 20 s longer than the forebody heating proe le. Bondline temperaturedata, although not useful forquantitative analysis of aerothermal heating, clearly showtheheatshield had adequate thickness margins for the actual entry.
125 citations
06 Jan 1997
103 citations
"Ablation and Thermal Response Progr..." refers methods in this paper
...Tables of B' (and hw) for ablative materials can be generated using the ACE or MAT codes.(8-9) In the equations above, Hr, ρeueCH1, αw, εw, λ and qrad are input quantities, and...
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