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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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Carbon-carbon materials and new oxidation resistant coating developments are discussed. Potential areas of application are highlighted. A short bibliography of selected references is included that describe carbon-carbon materials and related technology in detail.
An experimental program has been undertaken to better characterize the stress-strain characteristics of flexible material systems to support a NASA ground test program for inflatable decelerator material technology. A goal of the current study is to investigate experimental methods for the characterization of coated woven material stiffness. This type of experimental mechanics data would eventually be used to define the material inputs of fluid-structure interaction simulation models. The test methodologies chosen for this stress-strain characterization are presented along with the experimental results.
Examples of the contribution of microstructural analysis to the development of the Space Shuttle tile insulation system are presented, with photographic examples of the scanning electron microscope (SEM) investigations. After the basic thermal protection system materials had been selected, it was neccessary to analyze the mechanical responses of the combined materials; which included: (1) the polymer strain isolation pad (SIP), (2) the room temperature-vulcanizing silicone rubber bond, (RTV), and (3) rigid ceramic fiber reusable surface insulation (RSI). Microstructural analysis was used to provide information on deformation and fracture mechanisms, load transfer mechanisms, and structural alterations occurring before final failure. Both quantitative and qualitative information was obtained in the open, three-dimensional fibrous structures of the ceramic tiles by means of novel techniques of encapsulation and dissolution.
This presentation identifies the key materials with the highest payoff to advance the state-of-the-art in materials for turbomachinery. Current show stoppers for advancing the state of the art in materials and processes were identified. Technical issues associated with incorporating key materials into new systems are discussed. Opportunities, where they exist, are identified to overcome these technical challenges.
Multiscale micromechanics predictions have been made for the effective thermal conductivities and local thermal fields for a novel 3D woven composite thermal protection system material. The Multiscale Recursive Micromechanics approach, which enables micromechanics models to call other micromechanics models (or themselves recursively) to consider finer and finer length scales, has been employed. The multiscale model uses a recently developed version of the High-Fidelity Generalized Method of Cells micromechanics theory at each of three length scales. Results focus on the impact of the microstructural geometry representation at each length scale on the material’s effective thermal conductivity, along with the local thermal flux and temperature fields induced in the microstructures.
Multiscale micromechanics predictions have been made for the effective thermal conductivities and local thermal fields for a novel 3D woven composite thermal protection system material. The Multiscale Recursive Micromechanics approach, which enables micromechanics models to call other micromechanics models (or themselves recursively) to consider finer and finer length scales, has been employed. The multiscale model uses a recently developed version of the High-Fidelity Generalized Method of Cells micromechanics theory at each of three length scales. Results focus on the impact of the microstructural geometry representation at each length scale on the material’s effective thermal conductivity, along with the local thermal flux and temperature fields induced in the microstructures.
The Sheathed Miniature Aerothermal Reentry Thermocouple is a micro-miniature thermocouple for high temperature measurement in extreme environments. It is available for use in Thermal Protection System materials for ground testing and flight. This paper discusses the heritage, and design of the instrument. Experimental and analytical methods used to verify its performance and limitations are described.
Numerous unexpected operational issues relating to the abrasive nature of lunar dust, such as scratched visors and spacesuit pressure seal leaks, were encountered during the Apollo missions. To avoid reoccurrence of these unexpected detrimental equipment problems on future missions to the Moon, a series of two- and three-body abrasion tests were developed and conducted in order to begin rigorously characterizing the effect of lunar dust abrasiveness on candidate surface system materials. Two-body scratch tests were initially performed to examine fundamental interactions of a single particle on a flat surface. These simple and robust tests were used to establish standardized measurement techniques for quantifying controlled volumetric wear. Subsequent efforts described in the paper involved three-body abrasion testing designed to be more representative of actual lunar interactions. For these tests, a new tribotester was developed to expose samples to a variety of industrial abrasives and lunar simulants. The work discussed in this paper describes the three-body hardware setup consisting of a rotating rubber wheel that applies a load on a specimen as a loose abrasive is fed into the system. The test methodology is based on ASTM International (ASTM) B611, except it does not mix water with the abrasive. All tests were run under identical conditions. Abraded material specimens included poly(methyl methacrylate) (PMMA), hardened 1045 steel, 6061-T6 aluminum (Al) and 1018 steel. Abrasives included lunar mare simulant JSC- 1A-F (nominal size distribution), sieved JSC-1A-F (<25 m particle diameter), lunar highland simulant NU-LHT-2M, alumina (average diameter of 50 m used per ASTM G76), and silica (50/70 mesh used per ASTM G65). The measured mass loss from each specimen was converted using standard densities to determine total wear volume in cm3. Abrasion was dominated by the alumina and the simulants were only similar to the silica (i.e., sand) on the softer materials of aluminum and PMMA. The nominal JSC- 1A-F consistently showed more abrasion wear than the sieved version of the simulant. The lunar dust displayed abrasivity to all of the test materials, which are likely to be used in lunar landing equipment. Based on this test experience and pilot results obtained, recommendations are made for systematic abrasion testing of candidate materials intended for use in lunar exploration systems and in other environments with similar dust challenges.
The task of managing hydrogen entry into elevated temperature structural materials employed in turbomachinery is a critical engineering area for propulsion systems employing hydrogen or decomposable hydrocarbons as fuel. Extant structural materials, such as the Inconel series, are embrittled by the ingress of hydrogen in service, leading to a loss of endurance and general deterioration of load-bearing dependability. Although the development of hydrogen-insensitive material systems is an obvious engineering option, to date insensitive systems cannot meet the time-temperature-loading service extremes encountered. A short-term approach that is both feasible and technologically sound is the development and employment of hydrogen barrier coatings. The present project is concerned with developing, analyzing, and physically testing laminate composite hydrogen barrier systems, employing Inconel 718 as the structural material to be protected. Barrier systems will include all metallic, metallic-to-ceramic, and, eventually, metallic/ceramic composites as the lamellae. Since space propulsion implies repetitive engine firings without earth-based inspection and repair, coating durability will be closely examined, and testing regimes will include repetitive thermal cycling to simulate damage accumulation. The target accomplishments include: generation of actual hydrogen permeation data for metallic, ceramic-metallic, and hybrid metallic/ceramic composition barrier systems, practically none of which is currently extant; definition of physical damage modes imported to barrier systems due to thermal cycling, both transient temperature profiles and steady-state thermal mismatch stress states being examined as sources of damage; and computational models that incorporate general laminate schemes as described above, including manufacturing realities such as porosity, and whatever defects are introduced through service and characterized during the experimental programs.
• The flexible thermal protection system (FTPS) on NASA’s Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) vehicle will be instrumented with thermocouples (TCs) to measure the in-depth thermal response during entry into Earth’s atmosphere. Accurate flight temperature measurements are critical for verifying vehicle performance and reducing thermal model uncertainties. • The TC leads are routed from the measurement location to the data acquisition system within the same FTPS layer that they are monitoring the temperature. This approach eliminates the need to put holes in the FTPS layers. However, the insulated TC leads are exposed to high temperatures and large thermal gradients. • The baseline TCs were commercially available 30 AWG Type K TCs with a binder impregnated glass braid insulation. The glass braid is rated to a maximum continuous use temperature of 482°C. LOFTID’s heat pulse will be on the order of minutes and the maximum predicted temperature beneath the outermost FTPS layers is 1350°C.
All spacecraft require propulsion systems for thrust and maneuvering. Propulsion systems can be chemical, nuclear, electrical, cold gas or combinations thereof. Chemical propulsion has proven to be the most reliable technology since the deployment of launch vehicles. Performance, storability, and handling are three important aspects of liquid chemical propulsion. Bipropellant systems require a fuel and an oxidizer for propulsion, but monopropellants only require a fuel and a catalyst for propulsion and are therefore simpler and lighter. Hydrazine is the state of the art propellant for monopropellant systems, but has drawbacks because it is highly hazardous to human health, which requires extensive care in handling, complex ground ops due to safety and environmental considerations, and lengthy turnaround times for reusable spacecraft. All users of hydrazine monopropellant must contend with these issues and their associated costs. The development of a new monopropellant, intended to replace hydrazine, has been in progress for years. This project will apply advanced techniques to characterize the engineering properties of materials used in AF-M315E propulsion systems after propellant exposure. AF-M315E monopropellant has been selected HQ's Green Propellant Infusion Mission (GPIM) to replace toxic hydrazine for improved performance and reduce safety and health issues that will shorten reusable spacecraft turn-around time. In addition, this project will fundamentally strengthen JSC's core competency to evaluate, use and infuse liquid propellant systems.
The optical characteristics of surfaces on spacecraft are fundamental parameters in controlling its temperature. Passive thermal control coatings with designed solar absorptance and infrared emittance properties have been developed and been in use for some time. In this total space environment, the coating must be stable and maintain its desired optical properties for the course of the mission lifetime. The mission lifetimes are increasing and in our quest to save weight, newer substrates are being integrated which limit electrical grounding schemes. All of this has already added to the existing concerns about spacecraft charging and related spacecraft failures or operational failures. The concern is even greater for thermal control surfaces that are very large. One way of alleviating such concerns is to design new thermal control material systems (TCMS) that can help to mitigate charging via providing charge leakage paths. The object of this program was to develop two types of passive electrically conductive TCMS.
To contribute to and complement our understanding of the processes governing the formation,distribution, and evolution of primitive materials throughout the solar system, it will be critical toform connections between broad remote sensing spectroscopic surveys, laboratorymeasurements of analogs and samples delivered to Earth, and in situ measurements of thesurface composition on future primitive body missions. Recently, a laboratory prototypeemploying resonance two-step laser mass spectrometry [Getty et al., 2012] has been coupledto a cryogenic sample stage to enable measurements of analog samples that are relevant tothese fundamental questions. Analyses of mineral-aromatic mixtures and meteorite powderswill be presented. Our goals are twofold: (1) to conduct laboratory studies on solar systemanalog, meteoritic, and potentially returned samples to elucidate composition, and (2) toprovide a compact but capable analytical instrument for discovery-driven in situ interrogationof surface chemistry on a future mission, such as to a Trojan asteroid, comet, or icy moon.
Carbon preforms used in Thermal Protection System (TPS) materials are 80 to 90% porous, allowing for boundary layer and pyrolysis gases to flow through the porous regions. The bulk material properties such as permeability and hydraulic tortuosity factor affect the transport of the boundary layer gases. The use of Direct Simulation Monte Carlo along with the Klinkenberg permeability formulation allows us to compute the continuum permeability and Knudsen correction factor for flow in the transition regime. In this work, we have computed the permeability for two types of carbon preforms, namely, Morgan Felt and FiberForm, and assessed the effect of orientation on the permeability. Since both the materials are anisotropic, the permeability was found to depend on orientation, wherein, the materials are more permeable in the in-plane orientation than the through-thickness orientation. The through-thickness orientation was also more tortuous compared to the in-plane material orientation. Compared to Morgan Felt, FiberForm is less permeable, in both, through thickness and in-plane directions.
Tests were conducted at room temperature to determine the mechanical properties and behavior of materials used for the thermal protection system of the space shuttle. The materials investigated include the LI-900 RSI tiles, the RTV-560 adhesive and the .41 cm (.16 thick) strain isolator pad (SIP). Tensile and compression cyclic loading tests were conducted on the SIP material and stress-strain curves obtained for various proof loads and load cyclic conditioning. Ultimate tensile and shear tests were conducted on the RSI, RTV, and SIP materials. The SIP material exhibits highly nonlinear stress-strain behavior, increased tangent modulus and ultimate tensile strength with increased loading rate, and large short time load relaxation and moderate creep behavior. Proof and cyclic load conditioning of the SIP results in permanent deformation of the material, hysteresis effects, and much higher tensile tangent modulus values at large strains.
A standard Hastelloy-X honeycomb material and a pack aluminide coated honeycomb material were evaluated as to their performance as labyrinth seal materials for aircraft gas turbine engines. Consideration from published literature was given to the fluid sealing characteristics of two honeycomb materials in labyrinth seal applications, and their rub characteristics, erosion resistance, and oxidation resistance were evaluated. The increased temperature potential of the coated honeycomb material compared to the uncoated standard could be achieved without compromising the honeycomb material's rub tolerance, although there was some penalty in terms of reduced erosion resistance.
Metal hydrides are serious contenders for materials-based hydrogen storage to overcome constraints associated with compressed or liquefied H 2 . Their ultimate performance is usually evaluated using intrinsic material properties without considering a systems design perspective. An illustrative case with startling implications is (LiNH 2 +2LiH). Using models that simulate the storage system and associated fuel cell of a light-duty vehicle (LDV), the performance of the bulk hydrides is compared with a nanoscaled version in porous carbon (PC), (LiNH 2 +2LiH)@(6-nm PC). Using experimental material properties, the simulations show that (LiNH 2 +2LiH)@(6-nm PC) counterintuitively has higher usable gravimetric and volumetric capacities than the bulk counterpart on a system basis despite having lower capacities on a materials-only basis. Nanoscaling increases the thermal conductivity and lowers the desorption enthalpy, which consequently increases heat management efficiency. In a simulated drive cycle for fuel cell-powered LDV, the fuel cell is inoperable using bulk (LiNH 2 +2LiH) as the storage material but completes the drive cycle using the nanoscale material. Further, these results challenge the notion that nanoscaling incurs mass and volume penalties. Instead, the synergistic nanoporous host-hydride interaction can favorably modulate chemical and heat transfer properties. Moreover, a co-design approach considering application-specific tradeoffs is essential to accurately assess a material's potential for real-world hydrogen storage.
Spallation is the mechanical removal of small chunks of material gets removed typically due to high shear conditions of the flow field. This reduces the ability of the thermal protection system (TPS) material to protect the spacecraft as well as cause turbulence in the flow causing higher heating rates. In this work, we focus on the material removal through ablation and high shear flow within the boundary layer region of woven TPS material. Woven TPS (WTPS) material is the latest class of material developed by NASA, to be used within the next generation of space flights. They are complex interlocked weaves designed to create a rigid structure that is highly resistant to heat and can be easily designed and tailored for a wide variety of entry environments. Due to material removal resulting from chemical degradation, the structural integrity of TPS material is affected. Spallation occurs when this structurally compromised material is exposed to the high shear flow conditions within the boundary layer. In order to understand the spallation mechanism within WTPS material, we first perform the material removal simulations which occur primarily through oxidation to obtain the microstructure at various stages of degradation. These simulations are performed using the Porous Microstructure Analysis (PuMA) software developed at NASA Ames. The micro-structure geometry used within these simulations were generated artificially to be similar to the 3D weave architecture of MSR-EEV (Mars Sample Return - Earth Entry Vehicle). The various eroded TPS micro-structures are then subjected to the boundary layer flow conditions to obtain critical surface quantities which contribute to the structural failure mechanism such as heat flux, pressure, and shear stress. The direct simulation Monte Carlo (DSMC) methodology is used to perform these simulations in order to accurately capture the strong gradients within the high-temperature boundary layer flow over the intricate geometry of WTPS material. The boundary layer profile is directly taken from the Computational Fluid Dynamics (CFD) simulation and provided as boundary conditions to the DSMC inlet and outlet. Further, the variation of these properties as the microstructure undergoes changes due to oxidation is also investigated. Finally, these quantities are used as input in PuMA to understand the material expansion/compression and strain within the woven TPS geometry and help in developing a comprehensive spallation and structure failure model.