Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “graphitization”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 1,279 records · Page 71
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Graphite solubility and co-vesiculation in basalt-like melts at one-ATM

The identity and source of the vapor phase that caused lunar lava-fountaining and vesiculation in lunar basalts continues to be of interest because of its implications for the composition and state of the lunar interior and because of its implications for lunar resources. In light of the apparent near-absence of H2O on the Moon, it has been suggested that the vapor phase may be CO2-CO. This premise is supported by the presence of carbon on the surface of volcanic glass beads. However, although the rapid exsolution of CO2 from a melt during decompression may be consistent with firefountaining, it fails to provide a satisfying explanation for vesiculation in mare basalt where exsolution of the gas phase would more reasonably be related to cooling/crystallization at low pressure rather than decompression from high pressure. Also, geochemical trends in lunar volcanic glasses suggest that their source has an oxygen fugacity more reducing than the iron-wustite buffer, an oxygen fugacity that is inconsistent with presence of dissolved CO2-CO at depth. The results of experiments in which a vesicular 'basalt' is produced from a melt equilibrated with graphite and pure CO gas at one atmosphere pressure are reported. The vesiculation is apparently related to exsolution of CO or a CO species during cooling of the melt or growth of quench crystals. Additionally, particulate carbon dispersed through the quenched sample suggests that elemental carbon is either in solution in the melt prior to quenching or tends to go into suspension perhaps as colloid-like particles. These two observations may provide insight into the nature of fire-fountaining and vesiculation on the Moon.

Colson, R. O.↗
NASA NTRS1993

Graphite fiber textile preform/copper matrix composites

This project has the objective of exploring the use of graphite fiber textile preform/copper matrix composites in spacecraft heat transmitting and radiating components. The preforms are to be fabricated by braiding of tows and when infiltrated with copper will result in a 3-D reinforced, near net shape composite with improved specific properties such as lower density and higher stiffness. It is anticipated that the use of textile technology will result in a more robust preform and consequently better final composite; it is hard to anticipate what performance tradeoffs will result, and these will be explored through testing and characterization.

Filatovs, G. J.↗
NASA NTRS1994

Test and analysis of a stitched RFI graphite-epoxy panel with a fuel access door

A stitched RFI graphite-epoxy panel with a fuel access door was analyzed using a finite element analysis and loaded to failure in compression. The panel was initially 56-inches long and 36.75-inches wide and the oval access door was 18-inches long and 15-inches wide. The panel was impact damaged with impact energy of 100 ft-lb prior to compressive loading; however, no impact damage was detectable visually or by A-scan. The panel carried a failure load of 695,000 Ib and global failure strain of .00494 in/in. Analysis indicated the panel would fail due to collapse at a load of 688,100 Ib. The test data indicate that the maximum strain in a region near the access door was .0096 in/in and analysis indicates a local surface strain of .010 in/in at the panel's failure load. The panel did not fail through the impact damage, but instead failed through bolt holes for attachment of the access door in a region of high strain.

Jegley, Dawn C.↗
NASA NTRS1992

Properties of Graphite Fiber Reinforced Copper Matrix Composites for Space Power Applications

The thermal and mechanical properties of pitch-based graphite fiber reinforced copper matrix (Gr/Cu) composites usable for space applications such as radiator fins were investigated. Thermal conductivity was measured as a function of fiber volume fraction and architecture. Results showed for unidirectional P-100 Gr/Cu composites, the longitudinal thermal conductivity was nearly independent of fiber volume fraction. Transverse thermal conductivities (perpendicular to the fibers) were strongly affected by the fiber volume fraction with higher volume fractions resulting in lower thermal conductivities. The effect of architecture on thermal conductivity followed the cosine squared law for simple architectures. Insufficient data are available currently to model more complex architectures, but adding fibers in the direction of the heat flow increases the thermal conductivity as low conductivity plies are supplemented by high conductivity plies. Thermal expansion tests were conducted on the Gr fibers and Gr/Cu composites. The results show a considerable thermal expansion mismatch between the fibers and the Cu matrix. The longitudinal thermal expansion showed a strong dependence on the architecture of the Gr/Cu composites. The composites also show a thermal expansion hysteresis. The hysteresis was eliminated by an engineered interface. Mechanical testing concentrated on the dynamic modulus and strength of the composites. The dynamic modulus of the Gr/Cu composites was 305 GPa up to 400 C, a value equivalent to Be. The strengths of the composites were less than expected, but this is attributed to the poor bond across the interface between the Gr fibers and Cu matrix. Testing of composites with an engineered interface is expected to yield strengths nearer the values predicted by the rule of mixtures.

Ellis, David L.↗
NASA NTRS1993

Interstellar grains in primitive meteorites - Diamond, silicon carbide, and graphite

Primitive meteorites contain a few parts per million (ppm) of pristine interstellar grains that provide information on nuclear and chemical processes in stars. Their interstellar origin is proven by highly anomalous isotopic ratios, varying more than 1000-fold for elements such as C and N. Most grains isolated thus far are stable only under highly reducing conditions (C/O greater than 1), and apparently are 'stardust' formed in stellar atmospheres. Microdiamonds, of median size about 10 A, are most abundant (about 400-1800 ppm) but least understood. They contain anomalous noble gases including Xe-HL, which shows the signature of the r- and p-processes. Silicon carbide, of grain size 0.2-10 microns and abundance about 6 ppm, shows the signature of the s-process and apparently comes mainly from red giant carbon (AGB) stars of 1-3 solar masses. Some grains appear to be not less than 10 exp 9 a older than the solar system. Graphite spherules of grain size 0.8-7 microns and abundance less than 2 ppm contain highly anomalous C and noble gases, as well as large amounts of fossil Mg-26 from the decay of extinct Al-26. They seem to come from at least three sources, probably AGB stars, novae, and Wolf-Rayet stars.

Anders, Edward↗
NASA NTRS1994

Selection process for trade study: Graphite Composite Primary Structure (GCPS)

This TA 2 document describes the selection process that will be used to identify the most suitable structural configuration for an SSTO winged vehicle capable of delivering 25,000 lbs to a 220 nm circular orbit at 51.6 degree inclination. The most suitable unpressurized graphite composite structures and material selections is within this configuration and will be the prototype design for subsequent design and analysis and the basis for the design and fabrication of payload bay, wing, and thrust structure full scale test articles representing segments of the prototype structures. The selection process for this TA 2 trade study is the same as that for the TA 1 trade study. As the trade study progresses additional insight may result in modifications to the selection criteria within this process. Such modifications will result in an update of this document as appropriate.

Greenberg, H. S.↗
NASA NTRS1995

Structural Arrangement Trade Study. Volume 1: Reusable Hydrogen Composite Tank System (RHCTS) and Graphite Composite Primary Structures (GCPS). Executive summary

This volume is the first of a three volume set that discusses the structural arrangement trade study plan that will identify the most suitable configuration for an SSTO winged vehicle capable of delivering 25,000 lbs to a 220 nm circular orbit at 51.6 deg inclination. The Reusable Hydrogen Composite Tank System (RHCTS), and Graphite Composite Primary Structures most suitable for intertank, wing and thrust structures are identified. This executive summary presents the trade study process, the selection process, requirements used, analysis performed and data generated. Conclusions and recommendations are also presented.

Source record↗
NASA NTRS1995

Structural arrangement trade study. Volume 3: Reusable Hydrogen Composite Tank System (RHCTS) and Graphite Composite Primary Structures (GCPS). Addendum

This volume is the third of a 3 volume set that addresses the structural trade study plan that will identify the most suitable structural configuration for an SSTO winged vehicle capable of delivering 25,000 lbs to a 220 nm circular orbit at 51.6 deg inclination. The most suitable Reusable Hydrogen Composite Tank System (RHCTS), and Graphite Composite Tank System (GCPS) composite materials for intertank, wing and thrust structures are identified. Vehicle resizing charts, selection criteria and back-up charts, parametric costing approach and the finite element method analysis are discussed.

Source record↗
NASA NTRS1993

A study of structurally efficient graphite-thermoplastic trapezoidal-corrugation sandwich and semi-sandwich panels

The structural efficiency of compression-loaded trapezoidal-corrugation sandwich and semi-sandwich composite panels is studied to determine their weight savings potential. Sandwich panels with two identical face sheets and a trapezoidal corrugated core between them, and semi-sandwich panels with a corrugation attached to a single skin are considered. An optimization code is used to find the minimum weight designs for critical compressive load levels ranging from 3,000 to 24,000 lb/in. Graphite-thermoplastic panels based on the optimal minimum weight designs were fabricated and tested. A finite-element analysis of several test specimens was also conducted. The results of the optimization study, the finite-element analysis, and the experiments are presented.

Jegley, Dawn C.↗
NASA NTRS1995

Graphite-Fiber Heat Radiators

Heat radiators of proposed type feature thermally conductive fibers protruding from metallic surfaces to provide increased heat-dissipation surface areas. Free of leaks and more reliable than radiators incorporating heat pipes. Also lightweight and relatively inexpensive. Radial graphite fibers carry heat away from spherical shell and radiate heat into space. Radiators prove useful on Earth in special industrial and scientific applications involving dissipation of heat in vacuum or in relatively still air.

Phillips, Wayne M.↗
NASA NTRS1995

Composite Of Bisfuran/Bismaleimide With Graphite Fibers

Composite-material specimens made of graphite fibers and matrix composed of experimental copolymer of bismaleimide and bisfuran. Found to have mechanical and thermal properties comparable or superior to those of similar composite-material specimens. Candidates for use as lightweight structural components to replace metal components in high-temperature aircraft applications. Bismaleimides used widely for other purposes because of low cost and ease in fabrication. Mechanical properties equal or exceed those of composite made with commercial formulation.

Meador, Mary Ann B.↗
NASA NTRS1995

Analysis of selected compression splice joint locations in a graphite-epoxy transport wing stub box

Three critical compression splice joint locations in a stitched graphite-epoxy transport wing stub box have been analyzed to determine their expected structural performance. The wing box is representative of a section of a commercial transport wing box and was designed and constructed by McDonnell Douglas Aerospace Company as part of the NASA Advanced Composites Technology (ACT) program. The results of the finite element analyses of the splice joints are presented. The analysis results indicate that failure will not occur in the splice joint regions for loads less than the Design Ultimate Load of the wing box.

Jegley, Dawn C.↗
NASA NTRS1995

Magnetic properties of iron-bearing graphite fibers

Carbon fibers containing ferromagnetically-ordered iron or other transition metals could be used in a variety of lightweight magnetic composites. Intercalation of bulk graphite with CoCl2 or FeCl3, followed by reduction with butyl lithium, did indeed produce magnetic samples; however, the observed room temperature permeabilities (mu) were less than 2 G/Oe. Magnetic data are presented on carbon fibers containing large amounts of elemental iron, which were prepared by a new method. In this method Amoco P-75 fibers were intercalated with Br2 and I2 followed by fluorination, forming CF(0.75). This product was then intercalated with FeCl3. Subsequent heat treatments in oxidizing and reducing atmospheres converted most of the iron to the pure alpha phase. Room temperature permeabilities as large as 40 G/Oe are observed.

Ferrante, R.↗
NASA NTRS1995

The Dependence of the Change in the Coefficient of Thermal Expansion of Graphite Fiber Reinforced Polyimide IM7-K3B on Microcracking due to Thermal Cycling

Composite IM7-K3B was subjected to a simulated high speed aircraft thermal environment to determine the effects of microcracking on the change in CTE. IM7-K3B is a graphite fiber reinforced polyimide laminate, manufactured by Dupont. The lay-up for the material was (0.90((Sub 3)(Sub s))). The specimens were placed in a laser-interferometric dilatometer to obtain thermal expansion measurements and were then repeatedly cycled between -65 F and 350 F up to 1000 cycles. After cycling they were scanned for microcracks at a magnification of 400x. The material was expected not to crack and to have a near zero CTE. Some microcracking did occur in all specimens and extensive microcracking occurred in one specimen. Further testing is required to determine how closely the CTE and microcracking are related.

Stewart, Melissa C.↗
NASA NTRS1973

Ablation of ATJ Graphite at High Temperatures

Results of an extensive experimental program on the ablation of ATJ graphite in air at surface pressures of 0.3 to 4.4 atm and surface temperatures of 2570 to 4030 K are presented. The measured mass-loss rates are correlated with surface temperature, pressure, and effective nose radius. The results are compared with several equilibrium thermochemical ablation theories which differ basically in the chemical species that are considered and in the thermodynamic properties assigned to these species. All the theories predict about the same mass-loss rate in the diffusion-controlled oxidation regime and are in good agreement with the experimental results. At higher temperatures, however, the experimental and theoretical results do not agree. At temperatures above 3700 K, the experimental mass-loss rate becomes independent of pressure and an exponential function of temperature. A comparison of the high-temperature results with the theories indicates that all the theories, except the one based on the thermodynamic properties from the JANAF tables, are invalid because they overpredict the mass-loss rate. On the other hand, the theory based on the JANAF properties underpredicts the measured mass-loss rate by a factor of three at 4000 K. At least a portion of this difference between experiment and theory is attributed to particulate mass loss which is visually observed.

Thermochemical ablation↗
NASA NTRS1997

Fabrication and Evaluation of Graphite Fiber-Reinforced Polyimide Composite Tube Forms Using Modified Resin Transfer Molding

The techniques necessary for the fabrication of a complex three-dimensional tubular form using a PMR-type resin have been developed to allow for the construction of several tubes with good physical and mechanical properties. Employing established resin transfer molding practices, the relatively non-hazardous AMB-21 in acetone formulation was used to successfully impregnate four layers of AS4 braided graphite fiber preform previously loaded around an aluminum cylindrical core in an enclosed mold cavity. Using heat and vacuum, the solvent was evaporated to form a prepreg followed by a partial imidization and removal of condensation products. The aluminum core was replaced by a silicone rubber bladder and the cure cycle continued to the final stage of 550 F with a bladder internal pressure of 200 lbs/sq in while simultaneously applying a strong vacuum to the prepreg for removal of any additional imidization products. A combination of several modifications to the standard resin transfer molding methodology enabled the mold to 'breathe', allowing the imidization products a pathway for escape. AMB-21 resin was chosen because of the carcinogenic nature of the primary commercial polyimide PMR-15. The AMB-21 resin was formulated using commercially available monomers or monomer precursors and dissolved in a mixture of methyl alcohol and acetone. The viscosity of the resulting monomer solution was checked by use of a Brookfield rheometer and adjusted by adding acetone to an easily pumpable viscosity of about 600 cP. In addition, several types of chromatographic and thermal analyses were of the braids, and excess handling of the preforms broke some of the microscopic fibers, needlessly decreasing the strength of the finished part. In addition, three dimensional braided preforms with fibers along the length of the tube will be significantly stronger in tension than the braided preforms used in this study.

Exum, Daniel B.↗
NASA NTRS1998

Comparison of Graphite Fabric Reinforced PMR-15 and Avimid N Composites After Long Term Isothermal Aging at Various Temperatures

Extensive effort is currently being expended to demonstrate the feasibility of using high-performance, polymer-matrix composites as engine structural materials over the expected operating lifetime of the aircraft, which can extend from 18,000 to 30,000 hr. The goal is to develop light-weight, high-strength, and high-modulus materials for use in higher temperature sections of advanced 21 st century aircraft propulsion systems. To accomplish this goal, it is necessary to pursue the development of thermal and mechanical durability models for graphite-fiber-reinforced, polymer-matrix composites. Numerous investigations have been reported regarding the thermo-oxidative stability (TOS) of the polyimide PMR-15 (1-5). A significant amount of this work has been directed at edge and geometry effects, reinforcement fiber influences, and empirical modeling of high-temperature weight loss behavior. It is yet to be determined if the information obtained from the PMR-15 composite tests is applicable to other polyimide-matrix composites. The condensation-curing polymer Avimid N is another advanced composite material often considered for structural applications at high temperatures. Avimid N has better thermo-oxidative stability than PMR-15 (6), but the latter is more easily processed. The most comprehensive study of the thermo-oxidative stability of Avimid N neat resin and composites at 371 (infinity)C is found in Salin and Seferis (7). The purposes of the work described herein were to compare the thermal aging behavior of these two matrix polymers and to determine the reasons for and the consequences of the difference in thermal durability. These results might be of some use in improving future polymer development through the incorporation of the desirable characteristics of both polyimides.

Bowles, Kenneth J.↗
NASA NTRS1998

Adjusting Measured Weight Loss of Aged Graphite Fabric/PMR-15 Composites

The purposes of this study were to evaluate the growth of the surface damage layer in polymer matrix composites (PMC's) fabricated with graphite fabric reinforcement and to determine the effects of the cut-surface degradation on the overall thermo-oxidative (TOS) stability of these materials. Four important conclusions were made about the TOS behavior of T650-35/PNIR- 15 fabric-reinforced composites: (1) Three stages of composite weight loss were seen on the plot of weight loss versus aging time; (2) the depth of the cut-edge damage is related to the composite thickness; (3) the actual weight loss realized by a mechanical test specimen that has had all the aging-induced cut-edge damage removed during the preparation process is significantly less than the weight loss measured using specimens with a high percentage of cut edges exposed to the damaging environment; and (4) an extrapolation of a section of the weight loss curve can be used to obtain a more correct estimate of the actual weight loss after extended periods of aging at elevated temperatures.

Bowles, Kenneth J.↗