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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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Studies were performed to synthesize a novel class of bis (imide-amine) curing agents for epoxy matrix resins. Glass transition temperatures and char yield data of an epoxy cured with various bis (imide-amines) are presented. The room temperature and 350 F mechanical properties, and char yields of unidirectional graphite fiber laminates prepared with conventional epoxy and imide-modified epoxy resins are presented.
Addition of vanadium compounds prevents melting. For safety, structural plastic foam should turn into rigid char when it burns, without melting. Addition of small amounts of vanadium compounds to some epoxy resins promotes char formation.
Improved method of predicting and of nondestructively analyzing pyrolytic effects in, and thermomechanical properties of, polymers combines theories and techniques from several engineering and scientific disciplines. Evolved in effort to understand charring and erosion of carbon phenolic ablative material in nozzle of rocket engine. Also applicable to other polymers and particularly to nondestructive engineering analysis of specimens heated, charred, or burned.
This study demonstrates that coupling of a material thermal response code and a flow solver with nonequilibrium gas/surface interaction for simulation of charring carbon ablators can be performed using an implicit approach. The material thermal response code used in this study is the three-dimensional version of Fully Implicit Ablation and Thermal response program, which predicts charring material thermal response and shape change on hypersonic space vehicles. The flow code solves the reacting Navier-Stokes equations using Data Parallel Line Relaxation method. Coupling between the material response and flow codes is performed by solving the surface mass balance in flow solver and the surface energy balance in material response code. Thus, the material surface recession is predicted in flow code, and the surface temperature and pyrolysis gas injection rate are computed in material response code. It is demonstrated that the time-lagged explicit approach is sufficient for simulations at low surface heating conditions, in which the surface ablation rate is not a strong function of the surface temperature. At elevated surface heating conditions, the implicit approach has to be taken, because the carbon ablation rate becomes a stiff function of the surface temperature, and thus the explicit approach appears to be inappropriate resulting in severe numerical oscillations of predicted surface temperature. Implicit coupling for simulation of arc-jet models is performed, and the predictions are compared with measured data. Implicit coupling for trajectory based simulation of Stardust fore-body heat shield is also conducted. The predicted stagnation point total recession is compared with that predicted using the chemical equilibrium surface assumption
NASA’s Space Launch System (SLS) experienced deflagration of the acreage of the Core Stage base heat shield during the Green Run hot-fire test campaigns and Artemis I flight. The burning products from the base heat shield during Green Run hot-fire test operation led to concerns in aerothermal environments and thermal protection system (TPS) performance prior to Artemis I flight. Updated design thermal environments accounting for cork combustion were developed using Green Run measurements and assessed prior to Artemis I, resulting in increased TPS thickness for the base heat shield. This work documents a highly controlled ground test within the NASA Marshall Space Flight Center’s Propulsion Test Branch Hot Gas Facility (HGF) and investigates the P50 cork TPS response due to sensitivity changes in radiative heat rates and surface shear stress. Prior to testing at the HGF, most of the P50 cork panels had a thick layer of char due to the cork combustion process replicated in a thermal vacuum chamber. SLS ascent shear stress and heating rate profile were simulated within the HGF by running the combustion-driven wind tunnel at off-nominal conditions and Mach numbers below 1. This test data provided insight into aerothermal measurements from Artemis I flight as well as informing updated base heating design environments for Artemis II. An in-depth understanding of the TPS P50 cork char response and observations is presented for launch vehicle flight ascent environments.
Ablation of charring material by high velocity heated gas of different compositions - chemical reactions
Model polymer synthesis for use in char forming heat shields
Thermal conductance of chars from nylon - phenolic composites as function of graphitization, density, and pore spectra
Hemispherical spectral emission of ablation chars, carbon, and zirconium oxide formed during reentry
Analytical model for predicting transient one- dimensional thermal performance of charring- ablator heat-protection system exposed to hyperthermal environment
Radionuclide method for determining Apollo heat shield char-front recession during reentry
Transient thermoconductivity determinations on low density phenol nylon char
Measured directional reflectances of chars in ultraviolet, visible, and near infrared spectra
Effects of thermal-property variations on performance of charring ablator
Analytical study of diffusion-controlled char oxidation and effect on steady state ablation of plastic materials
Permeability measurements of low-density phenolic-nylon chars in gaseous nitrogen and helium
Optical reflectance measurements of pyrolytic graphite and phenolic nylon chars and potential application to heat shield requirements
Conductive heat transfer in porous carbonaceous chars of degraded phenolic nylon sample