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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS2016

Post-Flight Assessment of Avcoat Thermal Protection System for the Exploration Flight Test-1

On December 5, 2014 NASA conducted the first flight test of its next generation human-class Orion spacecraft. The flight was called the Exploration Flight Test -1 (EFT-1) which lasted for 4 hours and culminated into a re-entry trajectory at 9 km/s. This flight test of the 5-meter Orion Crew Module demonstrated various sub-systems including the Avcoat ablative thermal protection system (TPS) on the heat shield. The Avcoat TPS had been developed from the Apollo-era recipe with a few key modifications. The engineering for thermal sizing was supported by modeling, analysis, and ground tests in arc jet facilities. This paper will describe a postlfight analysis plan and present results from post-recovery inspections, data analysis from embedded sensors, TPS sample extraction and characterization in the laboratory. After the recovery of the vehicle, a full photographic survey and surface scans of the TPS were performed. The recovered vehicle showed physical evidence of flow disturbances, varying degrees of surface roughness, and excessive recession downstream of compression pads. The TPS recession was measured at more than 200 locations of interest on the Avcoat surface. The heat shield was then processed for sample extraction prior to TPS removal using the 7-Axis Milling machine at Marshall Space Flight Center. Around 182 rectangular TPS samples were extracted for subsequent analysis and investigation. The final paper will also present results of sample analysis. The planned investigation includes sidewall imaging, followed by image analysis to characterize TPS response by quantifying different layers in the char and pyrolysis zones. A full postmortem of the instrumentation and sensor ports will also be performed to confirm no adverse effects due to the sensors themselves. A subset of the samples will undergo structural testing and perform detailed characterization of any cracks and integrity of gore seams. Finally, the material will be characterized with layer-by-layer density measurements and SEM investigations to evaluate material morphology at microstructural level including identification of elements and compounds.

Aerothermodynamics↗
NASA NTRSDate not supplied

Coupling CFD and Material Response for Analysis of Mars Entry

In computing the response of an ablating thermal protection system during atmospheric entry, the aerothermal environment and material response are generally computed separately with a blowing correction term in the material response model to account for the blowing of char and pyrolysis gases [1]. In this work, we apply a coupled approach in which pyrolysis blowing gases, computed in the PATO material response code [2], are used with a blowing boundary condition in the DPLR hypersonic CFD code [3]. This leads to an iterative method in which blowing products from PATO are input into DPLR to update surface heating estimates. The full iterative method, with the addition of radiative heating estimates using the NEQAIR radiation solver [4], is shown in Fig. 1. The method is demonstrated on a sphere case with the environment and material properties based on the Mars Science Laboratory entry. Future work includes utilizing this method in computing full 3D material response during the Mars 2020 entry.

Heat Transfer↗
NASA NTRSDate not supplied

Coupling CFD and Material Response for Analysis of Mars Entry

In computing the response of an ablating thermal protection system during atmospheric entry, the aerothermal environment and material response are generally computed separately with a blowing correction term in the material response model to account for the blowing of char and pyrolysis gases. In this work, we apply a coupled approach in which pyrolysis blowing gases, computed in the PATO material response code, are used with a blowing boundary condition in the DPLR hypersonic CFD code. This leads to an iterative method in which blowing products from PATO are input into DPLR to update surface heating estimates. The full iterative method, with the addition of radiative heating estimates using the NEQAIR radiation solver, is shown in Fig. 1. The method is demonstrated on a sphere case with the environment and material properties based on the Mars Science Laboratory entry. Future work includes utilizing this method in computing full 3D material response during the Mars 2020 entry.

Ablation↗
NASA NTRSDate not supplied

Analysis of Mars 2020 Entry with Coupled Material Response and CFD

When computing the response of an ablating thermal protection system during entry, the aerothermal environment and material response are traditionally computed independently, with the introduction of a blowing correction term in the material response model to account for the outgassing of char and pyrolysis gases. This study presents an approach, where the pyrolysis blowing gases, calculated within the PATO material response code [1], are integrated into the DPLR hypersonic CFD code [2] via a blowing boundary condition. This methodology employs an iterative process, whereby the blown pyrolysis gas products from PATO are incorporated into DPLR, refining surface heating predictions. The NEQAIR code is used to compute radiative heating [3]. The method is applied in computing 3D material response of the Mars 2020 entry. When compared with uncoupled material response, the coupled simulations show a lower surface heat flux initially and a higher heat flux at peak heating as shown in Fig. 1. This is in agreement with previous work using a sphere case with environments from MSL [4].

Heat Transfer↗
NASA NTRSDate not supplied

Analysis of Mars 2020 Entry with Coupled Material Response and CFD

When computing the response of an ablating thermal protection system during entry, the aerothermal environment and material response are traditionally computed independently, with the introduction of a blowing correction term in the material response model to account for the outgassing of char and pyrolysis gases. This study presents an approach, where the pyrolysis blowing gases, calculated within the PATO material response code, are integrated into the DPLR hypersonic CFD code via a blowing boundary condition. This methodology employs an iterative process, whereby the blown pyrolysis gas products from PATO are incorporated into DPLR, refining surface heating predictions. The NEQAIR code is used to compute radiative heating. The method is applied in computing 3D material response of the Mars 2020 entry. When compared with uncoupled material response, the coupled simulations show a lower surface heat flux initially and a higher heat flux at peak heating as shown in Fig. 1. This is in agreement with previous work using a sphere case with environments from MSL.

Heat Transfer↗
NASA NTRS1972

A Numerical Analysis of the Transient Response of an Ablation System Including Effects of Thermal Nonequilibrium, Mass Transfer and Chemical Kinetics

The differential equations governing the transient response of a one-dimensional ablative thermal protection system undergoing stagnation ablation are derived. These equations are for thermal nonequilibrium effects between the pyrolysis gases and the char layer and kinetically controlled chemical reactions and mass transfer between the pyrolysis gases and the char layer. The boundary conditions are written for the particular case of stagnation heating with surface removal by oxidation or sublimation and pyrolysis of the uncharred layer occurring in a plane. The governing equations and boundary conditions are solved numerically using the modified implicit method (Crank-Nicolson method). Numerical results are compared with exact solutions for a number of simplified cases. The comparison is favorable in each instance.

Clark, R. K.↗
NASA NTRS2011

Development and Verification of the Charring, Ablating Thermal Protection Implicit System Simulator

The development and verification of the Charring Ablating Thermal Protection Implicit System Solver (CATPISS) is presented. This work concentrates on the derivation and verification of the stationary grid terms in the equations that govern three-dimensional heat and mass transfer for charring thermal protection systems including pyrolysis gas flow through the porous char layer. The governing equations are discretized according to the Galerkin finite element method (FEM) with first and second order fully implicit time integrators. The governing equations are fully coupled and are solved in parallel via Newton s method, while the linear system is solved via the Generalized Minimum Residual method (GMRES). Verification results from exact solutions and Method of Manufactured Solutions (MMS) are presented to show spatial and temporal orders of accuracy as well as nonlinear convergence rates.

Amar, Adam J.↗
NASA NTRS2010

Development and Verification of the Charring Ablating Thermal Protection Implicit System Solver

The development and verification of the Charring Ablating Thermal Protection Implicit System Solver is presented. This work concentrates on the derivation and verification of the stationary grid terms in the equations that govern three-dimensional heat and mass transfer for charring thermal protection systems including pyrolysis gas flow through the porous char layer. The governing equations are discretized according to the Galerkin finite element method with first and second order implicit time integrators. The governing equations are fully coupled and are solved in parallel via Newton's method, while the fully implicit linear system is solved with the Generalized Minimal Residual method. Verification results from exact solutions and the Method of Manufactured Solutions are presented to show spatial and temporal orders of accuracy as well as nonlinear convergence rates.

Amar, Adam J.↗
NASA NTRS1977

Assessment of relative flammability and thermochemical properties of some thermoplastic materials

Thermomechanical properties, flammability, oxygen index, relative toxicity of pyrolysis effluents, and char yields were studied for 12 advanced polymers which are candidates for use in aircraft interiors as decorative films, compression- and injection-molded parts and thermoplastic parts. Polymers sampled included polyphenylene sulfide, 9,9 bis (4-hydroxyphenol) fluorene polycarbonate-poly (dimethylsiloxane), polyether sulfone, polyvinyl fluoride and polyvinylidene fluoride. Availability of these samples, whether in commercial form or in test quantities, is specified. An estimate of relative fire resistance for the materials was obtained; the five polymers listed above were found to be the most fire resistant of the 12 sampled.

Kourtides, D. A.↗
NASA NTRS1994

Conversion of polymers of methyl- and vinylsilane to Si-C ceramics

Poly(methylsilane) and poly(vinylsilane) were synthesized using a titanocene catalyst, and their pyrolytic conversion to ceramics was followed using a combination of thermal analysis and infrared spectroscopy. The two polymers have distinctly different backbone structures, as determined by Si NMR; methylsilane polymerizes to a polysilane, while vinylsilane polymers have predominately polycarbosilane backbone, with some polysilane structure as well. The pyrolysis path and char yield were dependent primarily on backbone structure, with little influence of polymer molecular weight. The majority of the weight loss on conversion occurs below 650 degrees C, although bond rearrangement continues to 1400 degrees C. Poly(vinylsilane) produced a C-rich Si-C ceramic in which the carbon was dispersed on a sufficiently fine level to show resistance to oxidation on heating in air to 1400 degrees C.

Hurwitz, Frances I.↗
NASA NTRS1997

Experimental Observations of PMMA Spheres Burning at Reduced Gravity

Polymer combustion is a very complicated process which involves the coupling of gas-phase combustion, melting, pyrolysis, and possible charring of the condensed phase. Although only a few studies on the combustion of a spherically shaped polymer have been conducted, there is renewed interest in the subject. At the Third International Microgravity Combustion Workshop, we presented a preliminary experimental plan and proposed an apparatus to study the combustion of a PolyMethylMethAcrylate (PMMA) sphere at reduced gravity. In this paper, we describe the experimental hardware in detail, summarize our observations since the last workshop, and describe future studies. The main objective is to determine the burning rate of PMMA spheres at reduced gravity under different ambient oxygen concentrations and total pressures. The dependence of the burning rate on the initial sphere diameter will be examined. It is anticipated that the simple spherical geometry in conjunction with reduced gravity will facilitate an assessment of the effect of condensed phase behavior on polymer burning processes and can be used as a means to test polymer combustion models.

Yang, Ji-Ann C.↗
NASA NTRS1973

An analysis of a charring ablator with thermal nonequilibrium, chemical kinetics, and mass transfer

The differential equations governing the transient response of a one-dimensional ablative thermal protection system are presented for thermal nonequilibrium between the pyrolysis gases and the char layer and with finite rate chemical reactions occurring. The system consists of three layers (the char layer, the uncharred layer, and an optical insulation layer) with concentrated heat sinks at the back surface and between the second and third layers. The equations are solved numerically by using a modified implicit finite difference scheme to obtain solutions for the thickness of the charred and uncharred layers, surface recession and pyrolysis rates, solid temperatures, porosity profiles, and profiles of pyrolysis-gas temperature, pressure, composition, and flow rate. Good agreement is obtained between numerical results and exact solutions for a number of simplified cases. The complete numerical analysis is used to obtain solutions for an ablative system subjected to a constant heating environment. Effects of thermal, chemical, and mass transfer processes are shown.

Clark, R. K.↗