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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 NTRS2004

Pancam multispectral imaging results from the Opportunity Rover at Meridiani Planum

Panoramic Camera (Pancam) images from Meridiani Planum reveal a low-albedo, generally flat, and relatively rock-free surface. Within and around impact craters and fractures, laminated outcrop rocks with higher albedo are observed. Fine-grained materials include dark sand, bright ferric iron-rich dust, angular rock clasts, and millimeter-size spheroidal granules that are eroding out of the laminated rocks. Spectra of sand, clasts, and one dark plains rock are consistent with mafic silicates such as pyroxene and olivine. Spectra of both the spherules and the laminated outcrop materials indicate the presence of crystalline ferric oxides or oxyhydroxides. Atmospheric observations show a steady decline in dust opacity during the mission. Astronomical observations captured solar transits by Phobos and Deimos and time-lapse observations of sunsets.

Mars↗

NASA NTRS1997

Damage Initiation and Ultimate Tensile Strength of Scaled [0 deg n/90 deg n/0 deg n]sub T Graphite-Epoxy Coupons

Previous research on scaling effects in composite materials has demonstrated that the stress levels at first ply failure and ultimate failure of composite laminates are dependent on the size of the laminate. In particular, the thickness dimension has been shown to be the most influential parameter in strength scaling of composite coupons loaded in tension. Geometrically and constitutively scaled laminates exhibit decreasing strength with increasing specimen size, and the magnitude of the strength-size effect is a function of both material properties and laminate stacking sequence. Some of the commonly used failure criteria for composite materials such as maximum stress, maximum strain, and tensor polynomial (e.g., Tsai-Wu) cannot account for the strength-size effect. In this paper, three concepts are developed and evaluated for incorporating size dependency into failure criteria for composite materials. An experimental program of limited scope was performed to determine the first ply failure stress in scaled cross-ply laminates loaded in tension. Test specimens were fabricated of AS-4/3502 graphite-epoxy composite material with laminate stacking sequences of [0 deg n/90 deg n/o deg n]subT where n=1-6. Two experimental techniques were used to determine first ply failure, defined as a transverse matrix crack in the 90 deg ply: (1) step loading with dye penetrant x-ray of the specimen at each load interval, and (2) acoustic emission. The best correlation between first ply failure analysis and experimental data was obtained using a modified Weibull approach which incorporated the residual thermal stress and the outer ply constraint, as well as the ply thickness effect. Finally, a second set of experiments was performed to determine the tensile response and ultimate failure of the scaled cross-ply laminates. The results of these experiments indicated no influence of specimen size on tensile response or ultimate strength.

Jackson, Karen E.↗

NASA NTRS1995

A comparison of graphite/epoxy tape laminates and 2-D braided composites mechanical properties

A comparison of the mechanical properties of unidirectional composite tape laminates and of two-dimensional triaxially braided composite was conducted. The tape laminate layups were designed to match the percentage of axial fibers and the angle of the bias tows in the braided composite. The materials system used for the laminates is AS4/3501-6 which was chosen as the closest available match to As4/1895 used for the braids. The strength and stiffness properties measured here include tension, open-hole tension, filled-hole tension, compression and open-hole compression, all of these in both the longitudinal and transverse direction. Results show that the longitudinal modulus of both material forms is quite similar, but that the transverse modulus of the braids is lower. In terms of strength, the longitudinal unnotched strength of the braids is lower than that of the laminates, while the transverse strength is significantly lower. For both strength and stiffness, the crimp in the bias tows of the braid is probably the main cause for reduced properties. On the other hand, a very significant increase in open-hole and filled-hole tension strength was observed for the braids compared to the tape laminates. However, this was not observed in compression where all the braid properties are lower than for the laminates.

Minguet, Pierre J.↗

NASA NTRS1985

Fracture characteristics of angleplied laminates fabricated from overaged graphite/epoxy prepreg

A series of angleplied graphite/epoxy laminates was fabricated from overaged prepreg and tested in tension to investigate the effects of overaged or advanced cure material on the degradation of laminate strength. Results, which include fracture stresses, indicate a severe degradation in strength. In addition, the fracture surfaces and microstructural characteristics are distinctly unlike any features observed in previous tests of this prepreg and laminate configuration. Photographs of the surfaces and microstructures reveal flat morphologies consisting of alternate rows of fibers and hackles. These fracture surface characteristics are independent of the laminate configurations. The photomicrographs are presented and compared with data from similar studies to show the unique characteristics produced by the overage prepreg. Analytical studies produced results which agreed with those from the experimental investigations.

Ginty, C. A.↗

NASA NTRS1987

Fracture characteristics of angleplied laminates fabricated from overaged graphite/epoxy prepreg

Ginty, Carol A.↗

NASA NTRS1965

Fiber-base graphites.

Semicommercial fiber-base graphites including laminated composites as structural materials for applications to 6000 degrees F

STRUCTURAL MATERIAL↗

NASA NTRS1969

Self-lubricating gear

Self-lubricating gear, designed for long term operation in a vacuum at high, low, and ambient temperatures, is constructed of alternating layers of metal and a dry lubricant material, such as polytetrafluoroethylene, with a suitable reinforcing material bonded into a laminated composite unit, which is machined to form a standard gear.

Demorest, K. E.↗

NASA NTRS1971

Development status of reusable nonmetallic thermal protection

Development of reusable oxidation resistant carbon-carbon laminates and surface insulation materials for thermal protection systems

Greenshields, D.↗

NASA NTRS1978

Reproducibility of structural strength and stiffness for graphite-epoxy aircraft spoilers

Structural strength reproducibility of graphite epoxy composite spoilers for the Boeing 737 aircraft was evaluated by statically loading fifteen spoilers to failure at conditions simulating aerodynamic loads. Spoiler strength and stiffness data were statistically modeled using a two parameter Weibull distribution function. Shape parameter values calculated for the composite spoiler strength and stiffness were within the range of corresponding shape parameter values calculated for material property data of composite laminates. This agreement showed that reproducibility of full scale component structural properties was within the reproducibility range of data from material property tests.

Howell, W. E.↗

NASA NTRS1982

Fire-Retardant Epoxy Adhesives

Phosphorus-containing epoxy is fire-retardant and translucent. Intended as adhesive for laminated plastic sheets, new material bonds well to titanium dioxide-filled plastic film, which ordinarily shows little surface interaction with adhesives. Fire retardancy has been demonstrated, and smoke density is low enough to avoid smoke obscuration.

Bilow, N.↗

NASA NTRS1985

Effect of low-velocity or ballistic impact damage on the strength of thin composite and aluminum shear panels

Impact tests were conducted on shear panels fabricated from 6061-T6 aluminum and from woven fabric prepreg of Du Pont Kevlara fiber/epoxy resin and graphite fiber/epoxy resin. The shear panels consisted of three different composite laminates and one aluminum material configuration. Three panel aspect ratios were evaluated for each material configuration. Composite panels were impacted with a 1.27-cm (0.05-in) diameter aluminum sphere at low velocities of 46 m/sec (150 ft/sec) and 67 m/sec (220 ft/sec). Ballistic impact conditions consisted of a tumbled 0.50-caliber projectile impacting loaded composite and aluminum shear panels. The results of these tests indicate that ballistic threshold load (the lowest load which will result in immediate failure upon penetration by the projectile) varied between 0.44 and 0.61 of the average failure load of undamaged panels. The residual strengths of the panels after ballistic impact varied between 0.55 and 0.75 of the average failure strength of the undamaged panels. The low velocity impacts at 67 m/sec (220 ft/sec) caused a 15 to 20 percent reduction in strength, whereas the impacts at 46 m/sec (150 ft/sec) resulted in negligible strength loss. Good agreement was obtained between the experimental failure strengths and the predicted strength with the point stress failure criterion.

Farley, G. L.↗

NASA NTRS1991

Response of composite materials to low velocity impact

Orthotropic and quasiisotropic laminates clamped at the edges were impacted by an instrumented falling weight. Five matrix materials, comprising a baseline epoxy, two toughened thermosets, and amorphous and crystalline thermoplastics were studied. For each material, the projected damage area and residual compression strengths and strains were determined as a function of impact energy. Principal conclusions are: (1) incipient damage associated with a prominent load drop during the impact test seems to be decisive for residual properties; (2) subtle rate effects are present with some materials; and (3) results from the small scale tests parallel those in standard tests and may be adequate for screening purposes.

Srinivasan, K.↗

NASA NTRS1992

Removal of surface loop from stitched composites can improve compression and compression-after-impact strengths

Stitching through-the-thickness (TTT) of composite materials produces a surface loop of yarn between successive penetrations. The surface loop is pressed into the surface layers of the composite material during the curing of the laminate, kinking the in-plane fibers near the surface of the material. The compression strength and compression-after-impact (CAI) strengths of carbon-epoxy specimens were measured with and without the surface loop. Removal of the surface loop had no influence on failure mode or failure mechanism, but did significantly increase the compression and CAI strengths.

Farley, Gary L.↗

NASA NTRS1992

Thermal residual stresses in silicon-carbide/titanium (0/90) laminate

The current work formulated a micromechanical analysis of a cross-ply laminate and calculated the thermal residual stress in a very thick (0/90)(sub 2n) silicon-carbide/titanium laminate. Results were also shown for a unidirectional laminate of the same material. Discrete fiber-matrix models assuming a rectangular array of fibers with a fiber volume fraction of 32.5 percent and a three-dimensional, finite-element analysis were used. Significant differences in the trends and magnitudes for the fiber, matrix, and interface stresses were calculated for unidirectional and (0/90) models. Larger hoop stresses calculated for the (0/90) model indicate that it may be more susceptible to radial cracking when subjected to mechanical loading than the unidirectional model. The axial stresses in the matrix were calculated to be slightly larger for the (0/90) model. The compressive axial stresses in the fiber were significantly larger in the (0/90) model. The presence of the cross-ply in the (0/90) model reduced the constraint on the fiber, producing radial interface stresses that were less compressive, which could lead to earlier failure of the fiber-matrix interface.

Bigelow, C. A.↗

NASA NTRS1993

The effects of long-duration space exposure on the mechanical properties of some carbon-reinforced resin matrix composites

Long Duration Exposure Facility (LDEF) Experiment A0175 involved the non-instrumented exposure of seven carbon-fiber reinforced resin-matrix advanced composite panels contained in two trays - A7 and A1. These two trays were located, respectively, on the leading and trailing faces of LDEF, obliquely oriented to the RAM (Row 9) and WAKE (Row 3) directions. The identity and location of the seven panels, which consisted of six flat laminates of the following material systems are shown: carbon/epoxy (T300/934), carbon/bismaleimide (T300/F178), and carbon/polyimide (C6000/LARC-160 and C6000/PMR-15), plus one bonded honeycomb sandwich panel (T300/934 face sheets and Nomex core) patterned after the Space Shuttle payload bay door construction. These material systems were selected to represent a range of then-available matrix resins which, by virtue of their differing polymer chemistry, could conceivably exhibit differing susceptibility to the low-earth orbit (LEO) environment. The principal exposure conditions of the LDEF environment at these tray locations are shown. Noteworthy to some of the observations discussed is the four-orders-of magnitude difference in the atomic oxygen (AO) fluence, which made a shallow incidence angle (approximately 22 deg) to Tray A7, while Tray A1 on the trailing face was essentially shielded from AO exposure. This evaluation focused on determining the individual and relative suitability of a variety of resin-matrix composite systems for long-term space structural applications. This was accomplished primarily by measuring and comparing a range of engineering mechanical properties on over 300 test coupons sectioned from the flight panels and from identical control panels, and tested at ambient and elevated temperatures. This testing was supported by limited physical characterization, involving visual examination of flight panel surface features, measurements of weight loss and warpage, and examination for changes in internal integrity (micro cracking, delamination) by ultrasonic c-scan and polished cross-sections.

Vyhnal, Richard F.↗

NASA NTRS1992

Structural analysis and sizing of stiffened, metal matrix composite panels for hypersonic vehicles

The present method for strength and stability analyses of stiffened, fiber-reinforced composite panels to be used in hypersonic vehicle structures is of great generality, and can be linked with planar finite-element analysis (FEA). Nonlinear temperature and load-dependent material data for each laminate are used to 'build-up' the stiffened panel's membrane, bending, and membrane-bending coupling stiffness terms, as well as thermal coefficients. The resulting, FEA-solved thermomechanical forces and moments are used to calculate strain at any location in the panel; this allows an effective ply-by-ply orthotropic strength analysis to be conducted, together with orthotropic instability checks for each laminated segment of the cross-section.

Collier, Craig S.↗

NASA NTRS1993

Stiffness, thermal expansion, and thermal bending formulation of stiffened, fiber-reinforced composite panels

A method is presented for formulating stiffness terms and thermal coefficients of stiffened, fiber-reinforced composite panels. The method is robust enough to handle panels with general cross sectional shapes, including those which are unsymmetric and/or unbalanced. Nonlinear, temperature and load dependent constitutive material data of each laminate are used to 'build-up' the stiffened panel membrane, bending, and membrane-bending coupling stiffness terms and thermal coefficients. New thermal coefficients are introduced to quantify panel response from through-the-thickness temperature gradients. A technique of implementing this capability with a single plane of shell finite elements using the MSC/NASTRAN analysis program (FEA) is revealed that provides accurate solutions of entire airframes or engines with coarsely meshed models. An example of a composite, hat-stiffened panel is included to demonstrate errors that occur when an unsymmetric panel is symmetrically formulated as traditionally done. The erroneous results and the correct ones produced from this method are compared to analysis from discretely meshed three-dimensional FEA.

Collier, Craig S.↗

NASA NTRS1994

Computational simulation of composite structures with and without damage

A methodology is described which uses finite element analysis of various laminates to computationally simulate the effects of delamination damage initiation and growth on the structural behavior of laminated composite structures. The delamination area is expanded according to a set pattern. As the delamination area increases, how the structural response of the laminate changes with respect to buckling and strain energy release rate are investigated. Rules are presented for laminates of different configurations, materials and thickness. These results demonstrate that computational simulation methods can provide alternate methods to investigate the complex delamination damage mechanisms found in composite structures.

Wilt, Thomas F.↗