SEARCH · Search NASA
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.
Two techniques rapidly determine low-frequency resistivity of conductive epoxies and high-frequency dielectric properties of insulating epoxies. Conductive epoxy is molded in channels in plastic block. Four-point ohmmeter is used to apply current and sense voltage; it reads out resistance. Because mold has precise and stable dimensions, it produces accurate consistent measurements.
The feasibility of applying the in situ fiberization process to the fabrication of strain isolation pads (SIP) for the Space Shuttle and to the fabrication of graphite-epoxy composites was evaluated. The ISF process involves the formation of interconnected polymer fiber networks by agitation of dilute polymer solutions under controlled conditions. High temperature polymers suitable for SIP use were fiberized and a successful fiberization of polychloro trifluoroethylene, a relatively high melting polymer, was achieved. Attempts to fiberize polymers with greater thermal stability were unsuccessful, apparently due to characteristics caused by the presence of aromaticity in the backbone of such materials. Graphite-epoxy composites were fabricated by interconnecting two dimensional arrays of graphite fiber with polypropylene IS fibers with subsequent epoxy resin impregnation. Mechanical property tests were performed on laminated panels of this material to evaluate intralaminar and interlaminar shear strength, and thus fracture toughness. Test results were generally unpromising.
Annealing at temperatures below the glass transition temperature (sub-T/g/ annealing) in structural epoxy network glasses is found to affect long-term mechanical properties of symmetrically (+ or - 45 deg) reinforced 16-lamina graphite/epoxy composites. Post-cured specimens of Thornel 300 graphite/Narmco 5208 epoxy are sub-T(g) annealed at 140 C for about 10, 100, 1000, 10,000, and 100,000 min with a prior quenching above T(g). The ultimate tensile strength, strain-to-break, and static toughness of the composite materials are found to decrease as a function of sub-T(g) annealing time, while no weight loss is observed. Time-dependent changes in properties are explained by a physical aging concept based on free-volume decreases in the nonequilibrium glassy state of network epoxies, and results imply that changes in composite properties may occur with the service time.
An attempt is made to develop tough, moisture resistant, high char yield epoxy resins by means of novel bisimide amine (BIA) hardener curing agents and a state-of-the-art epoxy resin system. The BIAs are isolated as mixtures containing monomer, oligomer, and polymeric species, and then characterized by elemental analysis and high pressure liquid chromatography. The bisimide amine-cured epoxies (IMEs) were characterized with respect to moisture absorption, thermal properties, and physical and mechanical properties, as well as in the role of matrices in Celion 6000/IME composites. The relative toughness characteristics of each IME formulation was measured by the 10 deg off-axis tensile test, measuring the uniaxial tensile strength, shear strength, and shear-strain-to-failure of the composite systems.
The primary objective of this experiment is to accumulate the needed operational data associated with the exposure of graphite-polyimide and graphite-epoxy material to the environments of space. The experiment will be mounted in two 3-in.-deep peripheral trays. Graphite-polyimide specimens will occupy 1 1/3 trays and the graphite-epoxy specimens will occupy two-thirds of a tray. The experiment approach requires two matched sets of specimens with traceable records that are maintained for materials processing and specimen quality. After fabrication, one set of each test specimen will be sectioned and structurally tested to serve as a data baseline. After the flight, the other set of specimens will undergo extensive measurements of mechanical properties for comparison with the original data baseline. Structural testing of the graphite-polyimide specimens will provide strength and elastic data in tension, compression, and shear. Transverse tension microcracking and crack propagation will be evaluated by photomicroscopy. Structural testing of the graphite-epoxy specimens will include verification of laminate, core, adhesive, and fatigue properties as applied to the design and analysis of the payload bay door. Microcracking and crack propagation will also be analyzed by photomicroscopy.
A new, commercially available, trifunctional epoxy resin (tris-(hydroxyphenyl)-methane triglycidyl ether) was modified with synthetic rubber to increase the impact resistance of epoxy/graphite composites. These composites were reinforced with commercially available satin-weave carbon cloth using two formulations of epoxies (brominated and nonbrominated) containing various amounts of carboxy-terminated butadience acrylonitrile (CTBN) rubber that had been prereacted with epoxy resin. The impact resistance was determined by measuring the interlaminar shear strength of the composites after impact. The mechanical properties, such as flexural strength and modulus at room temperature and at 93 C, were also determined. Measurements were taken of the flammability and glass transition temperature (Tg); and a thermal-gravimetric analysis was made.
The thermal expansion of three epoxy-matrix composites, a polyimide-matrix composite and a borosilicate glass-matrix composite, each reinforced with continuous carbon fibers, has been measured and compared. The expansion of a composite with a rubber toughened epoxy-matrix and P75S carbon fibers was very different from the expansion of two different single phase epoxy-matrix composites with P75S fibers although all three had the same stacking sequence. Reasonable agreement was obtained between measured thermal-expansion data and results from classical laminate theory. The thermal expansion of a material may change markedly as a result of thermal cycling. Microdamage, induced by 250 cycles between -156 C and 121 C in the graphite/polyimide laminate, caused a 53 percent decrease in the coefficient of thermal expansion. The thermal expansion of the graphite/glass laminate was not changed by 100 thermal cycles from -129 C to 38 C; however, a residual strain of about 10 x 10 to the minus 6 power was measured for the laminate tested.
An experimental study is reported in which a nondestructive technique involving the use of a transparent fiberglass-epoxy composite birefringent material has been used to investigate compression failure mechanisms in graphite-epoxy laminates. It is shown that the birefringency and transparency of the fiberglass-epoxy material permits regions of high stress to be located and the mechanisms of local failure propagation to be identified within the laminate. The material may also be useful for studying stress fields and for identifying failure initiation and propagation mechanisms in a wide variety of composite-structure problems.
The effect of epoxy-size and degree of cure on the interfacial bonding of an epoxy-amine-graphite fiber composite system is examined. The role of the fiber-resin interface in determining the overall mechanical properties of composites is poorly understood. A good interfacial adhesive bond is required to achieve maximum stress transfer to the fibers in composites, but at the same time some form of energy absorbing interfacial interaction is needed to achieve high fracture toughening. The incompatibility of these two processes makes it important to understand the nature and basic factors involved at the fiber-resin interface as stress is applied. The mechanical properties including interlaminar shear values for graphite fiber-resin composites are low compared to glass and boron-resin composites. These differences have been attributed to poor fiber-matrix adhesion. Graphite fibers are commonly subjected to post-treatments including application of organic sizing in order to improve their compatibility with the resin matrix and to protect the fiber tow from damage during processing and lay-up. In such processes, sized graphite fiber tow is impregnated with epoxy resin and then layed-up i nto the appropriate configuration. Following an extended ambient temperature cure, the graphite-resin composite structure is cured at elevated temperature using a programmed temperature sequence to cure and then cool the product.
Studies were performed to synthesize new ether modified, flexibilized aromatic diamine hardeners for curing epoxy resins. The effect of moisture absorption on the glass transition temperatures of a tetraglycidyl epoxy, MY 720, cured with flexibilized hardeners and a conventional aromatic diamine was studied. Unidirectional composites, using epoxy-sized Celion 6000 graphite fiber as the reinforcement, were fabricated. The room temperature and 300 F mechanical properties of the composites, before and after moisture exposure, were determined. The Mode I interlaminar fracture toughness of the composites was characterized using a double cantilever beam technique to calculate the critical strain energy release rate.
The effects of the space environment on the engineering properties and chemistry of a chemically modified T300/934 graphite-epoxy composite system are characterized. The material was subjected to 1.0 x 10 to the 10th power rads of 1.0 MeV electron irradiation under vacuum to simulate 30 years in geosynchronous earth orbit. Monotonic tension tests were performed at room temperature (75 F/24 C) and elevated temperature (250 F/121 C) on 4-ply unidirectional laminates. From these tests, inplane engineering and strength properties (E sub 1, E sub 2, Nu sub 12, G sub 12, X sub T, Y sub T) were determined. Cyclic tests were also performed to characterize energy dissipation changes due to irradiation and elevated temperature. Large diameter graphite fibers were tested to determine the effects of radiation on their stiffness and strength. No significant changes were observed. Dynamic-mechanical analysis demonstrated that the glass transition temperature was reduced by 50 F(28 C) after irradiation. Thermomechanical analysis showed the occurrence of volatile products generated upon heating of the irradiated material. The chemical modification of the epoxy did not aid in producing a material which was more radiation resistant than the standard T300/934 graphite-epoxy system. Irradiation was found to cause crosslinking and chain scission in the polymer. The latter produced low molecular weight products which plasticize the material at elevated temperatures and cause apparent material stiffening at low stresses at room temperature.
This investigation of composite material properties utilized T300/934 graphite-epoxy that was subjected to 1.0 MeV electron radiation for a total dose of 1.0 x 10 to the 10th rads at a rate of 5.0 x 10 to the 7th rads/hour, simulating a worst-case exposure equivalent to 30 years in space. Mechanical testing was performed on 4-ply unidirectional laminates over the temperature range of -250 F (116 K) to +250 F (394 K). In-plane elastic tensile and shear properties as well as strength were obtained. The results show that electron radiation degrades the epoxy matrix and produces products that volatilize at the temperatures considered. These degradation products plasticize the epoxy at elevated temperatures and embrittle it at low temperatures, thereby altering the mechanical properties of the composite.
The thermal expansion of continuous carbon-fiber reinforced composites with epoxy-, polyimide-, and borosilicate glass-matrices has been measured and compared. The expansion of a rubber-toughened epoxy-matrix/P75S carbon-fiber composite was very different from the expansion of two different single-phase epoxy-matrix/P75S composites, although all three had the same stacking sequence. Reasonable agreement was obtained between measured thermal expansion data and results from classical laminate theory. Microdamage in the graphite/polyimide laminate, induced by 250 cycles between -156 and 121 C, caused a 53 percent decrease in the coefficient of thermal expansion. The thermal expansion of the graphite/glass laminate was not changed after 100 thermal cycles from -129 to 38 C; however, a residual strain of about 10 x 10 to the -6 was observed for the laminate tested.
Fiberglass-epoxy composites are considered for use as structural members for the mast of the space station solar array panel. The low Earth orbital environment in which space station is to operate is composed mainly of atomic oxygen, which has been shown to cause erosion of many organic materials and some metals. Ground based testing in a plasma asher was performed to determine the extent of degradation of fiberglass-epoxy composites when exposed to a simulated atomic oxygen environment. During exposure, the epoxy at the surface of the composite was oxidized, exposing individual glass fibers which could easily be removed. Several methods of protecting the composite were evaluated in an atomic oxygen environment and with thermal cycling and flexing. The protection techniques evaluated to date include an aluminum braid covering, an indium-tin eutectic and a silicone based paint. The open aluminum braid offered little protection while the CV-1144 coating offered some initial protection against atomic oxygen, but appears to develop cracks which accelerate degradation when flexed. Coatings such as the In-Sn eutectic may provide adequate protection by containing the glass fibers even though mass loss still occurs.
Postcured specimens of Thornell 300 carbon-fiber/Fiberite 934 epoxy as well as Fiberite 934 epoxy resin were quenched from above the glass transition temperature and annealed at 140, 110, or 80 C for times up to 100,000 min. No weight gain or loss was observed during annealing at these temperatures, but significant time-dependent variations were observed in density, modulus, hardness, damping, moisture absorption ability, and thermal expansion characteristics. The relations between the excess thermodynamic properties of epoxy glasses and their time-dependent behavior are discussed. The discussion is illustrated by results obtained by differential scanning calorimetry, dynamic mechanical analysis, tensile testing, solid state nuclear magnetic resonance spectroscopy, and small-angle X-ray scattering.
The extent of degradation of fiberglass-epoxy composite masts of the Space Station solar array panel, when these are exposed to atomic oxygen environment of the low-earth orbit, was investigated in ground testing of fiberglass-epoxy composites in an RF plasma asher. In addition, several methods of protecting the composite structures were evaluated, including an aluminum braid covering, an In-Sn eutectic, and a silicone based paint. It was found that, during exposure, the epoxy at the surface of the composite was oxidized, exposing individual glass fibers which could easily be removed. The results of mass measurements and SEM examination carried out after thermal cycling and flexing of exposed composite samples indicated that coatings such as In-Sn eutectic may provide adequate protection by containing the glass fibers, even though mass loss still occurs.
High-reflectance accurate-contour mirrors are needed for solar dynamic space power systems. Graphite fiber epoxy composites are attractive candidates for such applications owing to their high modulus, near-zero coefficient of thermal expansion, and low mass. However, mirrors prepared from graphite fiber epoxy composite substrates often exhibit print-through, a distortion of the surface, which causes a loss in solar specular reflectance. Efforts to develop mirror substrates without print-through distortion require a means of quantifying print-through. Methods have been developed to quantify the degree of print-through in graphite fiber epoxy composite specimens using surface profilometry.
An experimental investigation of the compression behavior of laminated specimens made from graphite-epoxy tape, graphite-thermoplastic tape and graphite-thermoplastic fabric was conducted. Specimens with five different stacking sequences were loaded to failure in uniaxial compression. Some of the specimens had central circular holes with diameters up to 65 percent of the specimen width. Other specimens were subjected to low speed impact with impact energies up to 35 J prior to compressive loading. This investigation indicates that graphite-thermoplastic specimens with holes have up to 15 percent lower failure stresses and strains than graphite-epoxy specimens with the same stacking sequence and hole size. However, graphite-thermoplastic specimens subjected to low speed impact have up to 15 percent higher failure stresses and strains than graphite-epoxy specimens with the same stacking sequence and impact energy. Compression tests of graphite-thermoplastic specimens constructed of unidirectional tape and fabric indicate that the material form has little effect on failure strains in specimens with holes or low speed impact damage.