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NASA NTRS1985

Characterization of Polyimide Matrix Resins and Prepregs

Graphite/polyimide composite materials are attractive candidates for a wide range of aerospace applications. They have many of the virtues of graphite/epoxies, i.e., high specific strengths and stiffness, and also outstanding thermal/oxidative stability. Yet they are not widely used in the aerospace industry due to problems of procesability. By their nature, modern addition polyimide (PI) resins and prepregs are more complex than epoxies; the key to processing lies in characterizing and understanding the materials. Chemical and rheological characterizations are carried out on several addition polyimide resins and graphite reinforced prepregs, including those based on PMR-15, LARC 160 (AP 22), LARC 160 (Curithane 103) and V378A. The use of a high range torque transducer with a Rheometrics mechanical spectrometer allows rheological data to be generated on prepreg materials as well as neat resins. The use of prepreg samples instead of neat resins eliminates the need for preimidization of the samples and the data correlates well with processing behavior found in the shop. Rheological characterization of the resins and prepregs finds significant differences not readily detected by conventional chemical characterization techniques.

Maximovich, M. G.↗

NASA NTRS1986

Fractography of unidirectional graphite-epoxy as a function of moisture, temperature and specimen quality

Optical microscopy and SEM have been used to examine the tensile failure surfaces of (0-deg)8 T300/5208 graphite-epoxy specimens, and fractography is employed to determine how moisture content and temperature, together with specimen preparation, affect failure modes. A low energy failure morphology is noted in defective specimens; specimens made from nondefective prepregs appeared to exhibit a decrease in flaw sensitivity and increasing strength with either temperature or moisture, although moisture also seemed to increase interfacial debonding between filament and matrix. The combination of temperature and moisture degraded performance by increasing interfacial debonding, and rendering the epoxy more prone to fracture.

Clements, L. L.↗

NASA NTRS1994

Micromechanical Modeling of the Thermal Expansion of Graphite/copper Composites with Nonuniform Microstructure

Two micromechanical models were developed to investigate the thermal expansion of graphite/copper (Gr/Cu) composites. The models incorporate the effects of temperature-dependent material properties, matrix inelasticity, initial residual stresses due to processing history, and nonuniform fiber distribution. The first model is based on the multiple concentric cylinder geometry, with each cylinder treated as a two-phase composite with a characteristic fiber volume fractions. By altering the fiber volume fraction of the individual cylinders, unidirectional composites with radially nonuniform fiber distributions can be investigated using this model. The second model is based on the inelastic lamination theory. By varying the fiber content in the individual laminae, composites with nonuniform fiber distribution in the thickness direction can be investigated. In both models, the properties of the individual regions (cylinders or laminae) are calculated using the method of cells micromechanical model. Classical incremental plasticity theory is used to model the inelastic response of the copper matrix at the microlevel. The models were used to characterize the effects of nonuniform fiber distribution on the thermal expansion of Gr/Cu. These effects were compared to the effects of matrix plasticity, choice of stress-free temperature, and slight fiber misalignment. It was found that the radially nonuniform fiber distribution has little effect on the thermal expansion of Gr/Cu but could become significant for composites with large fiber-matrix transverse CTE and Young's modulus mismatch. The effect of nonuniform fiber distribution in the through-thickness direction of a laminate was more significant, but only approached that of the stress-free temperature for the most extreme cases that include large amounts of bending. Subsequent comparison with experimental thermal expansion data indicated the need for more accurate characterization of the graphite fiber thermomechanical properties. Correlation with cyclic data revealed the presence of a mechanism not considered in the developed models. The predicted response did, however, exhibit ratcheting behavior that has been observed experimentally in Gr/Cu. Finally, simulation of the actual fiber distribution of particular specimens had little effect on the predicted thermal expansion.

Bednarcyk, Brett A.↗

NASA NTRS1994

Casting of weldable graphite/magnesium metal matrix composites with built-in metallic inserts

Technology innovations directed at the advanced development of a potentially low cost and weldable graphite/magnesium metal matrix composites (MMC) through near net shape pressure casting are described. These MMC components uniquely have built-in metallic inserts to provide an innovative approach for joining or connecting other MMC components through conventional joining techniques such as welding, brazing, mechanical fasteners, etc. Moreover, the metallic inserts trapped within the MMC components can be made to transfer the imposed load efficiently to the continuous graphite fiber reinforcement thus producing stronger, stiffer, and more reliable MMC components. The use of low pressure near net shape casting is economical compared to other MMC fabrication processes. These castable and potentially weldable MMC components can provide great payoffs in terms of high strength, high stiffness, low thermal expansion, lightweight, and easily joinable MMC components for several future NASA space structural, industrial, and commercial applications.

Lee, Jonathan A.↗

NASA NTRS2004

Failure of Non-Circular Composite Cylinders

In this study, a progressive failure analysis is used to investigate leakage in internally pressurized non-circular composite cylinders. This type of approach accounts for the localized loss of stiffness when material failure occurs at some location in a structure by degrading the local material elastic properties by a certain factor. The manner in which this degradation of material properties takes place depends on the failure modes, which are determined by the application of a failure criterion. The finite-element code STAGS, which has the capability to perform progressive failure analysis using different degradation schemes and failure criteria, is utilized to analyze laboratory scale, graphite-epoxy, elliptical cylinders with quasi-isotropic, circumferentially-stiff, and axially-stiff material orthotropies. The results are divided into two parts. The first part shows that leakage, which is assumed to develop if there is material failure in every layer at some axial and circumferential location within the cylinder, does not occur without failure of fibers. Moreover before fibers begin to fail, only matrix tensile failures, or matrix cracking, takes place, and at least one layer in all three cylinders studied remain uncracked, preventing the formation of a leakage path. That determination is corroborated by the use of different degradation schemes and various failure criteria. Among the degradation schemes investigated are the degradation of different engineering properties, the use of various degradation factors, the recursive or non-recursive degradation of the engineering properties, and the degradation of material properties using different computational approaches. The failure criteria used in the analysis include the noninteractive maximum stress criterion and the interactive Hashin and Tsai-Wu criteria. The second part of the results shows that leakage occurs due to a combination of matrix tensile and compressive, fiber tensile and compressive, and inplane shear failure modes in all three cylinders. Leakage develops after a relatively low amount of fiber damage, at about the same pressure for three material orthotropies, and at approximately the same location.

Hyer, M. W.↗

NASA NTRS1977

Lamination residual strains and stresses in hybrid laminates

An investigation is conducted of the effects of hybridization on the magnitude of lamination residual stresses. Eight-ply graphite/Kevlar 49/epoxy and graphite/S-glass/epoxy laminates were studied. The same matrix resin was selected for all basic materials to ensure compatibility and uniform curing of the various plies. The specimens, with inserted strain gages and thermocouples, were subjected to curing and postcuring cycles in an autoclave. Subsequently, the specimens were subjected to a thermal cycle from room temperature to 444 K and down to room temperature. It was found that hydridizing reduces apparently residual strains and stresses in the graphite plies. However, these strains were not affected much by the type and degree of hybridization.

Daniel, I. M.↗

NASA NTRS1983

Interlaminar shear properties of graphite fiber, high-performance resin composites

Short beam testing was used to determine the shear properties of laminates consisting of T-300 and Celion 3000 and 6000 graphite fibers, in epoxy, hot melt and solvent bismaleimide, polyimide and polystyrylpyridine (PSP). Epoxy, composites showed the highest interlaminar shear strength, with values for all other resins being substantially lower. The dependence of interlaminar shear properties on the fiber-resin interfacial bond and on resin wetting characteristics and mechanical properties is investigated, and it is determined that the lower shear strength of the tested composites, by comparison with epoxy resin matrix composites, is due to their correspondingly lower interfacial bond strengths. An investigation of the effect of the wettability of carbon fiber tow on shear strength shows wetting variations among resins that are too small to account for the large shear strength property differences observed.

Needles, H. L.↗

NASA NTRS1985

Fracture morphology of 2-D carbon-carbon composition

Out-of-plane tensile tests of a woven fabric carbon-carbon composite were performed in a scanning electron microscope equipped with a tensile stage and a videotape recording system. The composite was prepared from T-300 8-harness satin graphite fabric and a phenolic resin. The (0/90/0/90/0 sub 1/2) sub 2 laminate, with a Theta describing the orientation of the warp fibers of the fabric, was cured at 160 C and pyrolized at 871 C. This was followed by four cycles of resin impregnation, curing, and pyrolysis. A micrograph of the cross section of the composite is presented. Inspection of the specimen fracture surface revealed that the filaments had no residual matrix bonded to them. Further inspection revealed that the fracture was interlaminar in nature. Failure occurred where filaments of adjacent plies had the same orientation. Thus it is postulated that improvement in transverse tensile strength of 2-D carbon-carbon depends on the improvement of the filament-matrix bond strength.

Avery, W. B.↗

NASA NTRS1988

Advanced composite materials for precision segmented reflectors

The objective in the NASA Precision Segmented Reflector (PSR) project is to develop new composite material concepts for highly stable and durable reflectors with precision surfaces. The project focuses on alternate material concepts such as the development of new low coefficient of thermal expansion resins as matrices for graphite fiber reinforced composites, quartz fiber reinforced epoxies, and graphite reinforced glass. Low residual stress fabrication methods will be developed. When coupon specimens of these new material concepts have demonstrated the required surface accuracies and resistance to thermal distortion and microcracking, reflector panels will be fabricated and tested in simulated space environments. An important part of the program is the analytical modeling of environmental stability of these new composite materials concepts through constitutive equation development, modeling of microdamage in the composite matrix, and prediction of long term stability (including viscoelasticity). These analyses include both closed form and finite element solutions at the micro and macro levels.

Stein, Bland A.↗

NASA NTRS1986

Fatigue response of notched laminates subjected to tension-compression cyclic loads

The fatigue response of a ((0/45/90/-45)(sub s))(sub 4) T300-5208 graphite-epoxy laminate with a drilled center-hole subjected to various components of tensile and compressive cyclic loads was investigated. Damage evaluation techniques such as stiffness monitoring, penetrant-enhanced X-ray radiography, C-scan, laminate deply and residual strength measurement were used to establish the mechanisms of damage development as well as the effect of such damage on the laminate strength, stiffness and life. Damage modes consisted of transverse matrix cracks, initiating at the hole, in all plies, followed by delamination between plies of different orientation. A characteristic stiffness repsonse during cyclic loading at two load levels was identified and utilized a more reliable indicator of material and residual properties than accumulated cycles. For the load ratios of tension-compression loading, residual tensile strength increased significantly above the virgin strength early in the fatigue life and remained approximately constant to near the end of life. A technique developed for predicting delamination initiation sites along the hole boundary correlated well with experimental evidence.

Bakis, C. E.↗

NASA NTRS1995

Flexure fatigue testing of 90 deg graphite/epoxy composites

A great deal of research has been performed characterizing the in-plane fiber-dominated properties, under both static and fatigue loading, of advanced composite materials. To the author's knowledge, no study has been performed to date investigating fatigue characteristics in the transverse direction. This information is important in the design of bonded composite airframe structure where repeated, cyclic out-of-plane bending may occur. Recent tests characterizing skin/stringer debond failures in reinforced composite panels where the dominant loading in the skin is flexure along the edge of the frame indicate failure initiated either in the skin or else the flange, near the flange tip. When failure initiated in the skin, transverse matrix cracks formed in the surface skin ply closest to the flange and either initiated delaminations or created matrix cracks in the next lower ply, which in turn initiated delaminations. When failure initiated in the flanges, transverse cracks formed in the flange angle ply closest to the skin and initiated delamination. In no configuration did failure propagate through the adhesive bond layer. For the examined skin/flange configurations, the maximum transverse tension stress at failure correlates very well with the transverse tension strength of the composites. Transverse tension strength (static) data of graphite epoxy composites have been shown to vary with the volume of material stressed. As the volume of material stressed increased, the strength decreased. A volumetric scaling law based on Weibull statistics can be used to predict the transverse strength measurements. The volume dependence reflects the presence of inherent flaws in the microstructure of the lamina. A similar approach may be taken to determine a volume scale effect on the transverse tension fatigue behavior of graphite/epoxy composites. The objective of this work is to generate transverse tension strength and fatigue S-N characteristics for composite materials using 3-point flexure tests of 90 deg graphite/epoxy specimens. Investigations will include the volume scale effect as well as frequency and span-to-thickness ratio effects. Prior to the start of the experimental study, an analytical study using finite element modeling will be performed to investigate the span-to-thickness effect. The ratio of transverse flexure stress to shear stress will be monitored and its values predicted by the FEM analysis compared with the value obtained using a 'strength of materials' based approach.

Peck, Ann Nancy W.↗

NASA NTRS1982

Effects of thermal cycling on residual mechanical properties of C6000/PMR-15 graphite polyimide

An experimental study to determine the effects of thermal cycling on residual mechanical properties of a graphite/polyimide composite has been conducted. Interlaminar shear, flexure and compression strengths were measured at room temperature and 316 C on unidirectional and quasi-isotropic laminates subjected to 150 and 500 thermal cycles between -156 C and 316 C. Thermal cycling produced transverse microcracks and delaminations in the laminates. The linear density of the microcracks (35/cm) was about the same after 150 and 500 cycles. Although the crack density did not change from 150 to 500 cycles, crack growth did occur with intralamina cracks growing into adjacent lamina. The matrix controlled properties of the quasi-isotropic laminate at both room temperature and 316 C were significantly reduced by thermal cycling. However, the properties of the unidirectional laminate were significantly reduced only at 316 C.

Tompkins, S. S.↗

NASA NTRS1982

Failure characteristics of graphite-epoxy structural components loaded in compression

Failure characteristics of compressively loaded graphite-epoxy components are described. Experimental results for both strength-critical laminates and structural components with postbuckling strength are presented. Effects of low speed impact damage and circular holes on compressive strength are discussed. Delamination and shear crippling failure mechanisms that limit the performance of strength-critical laminates are described. Transverse shear and skin-stiffener separation failure mechanisms that limit the performance of components with postbuckling strength are also described. The influence of matrix properties on compressive strength improvements for impact damaged laminates is discussed. Experimental data and results from a failure analysis for strength-critical laminates with cutouts are discussed and compared with impact damage results. Typical postbuckling test results are compared with analytical predictions.

Starnes, J. H., Jr.↗

NASA NTRS1983

Failure characteristics of graphite-epoxy structural components loaded in compression

Starnes, J. H., Jr.↗

NASA NTRS2013

Next-Generation MKIII Lightweight HUT/Hatch Assembly

The MK III (H-1) carbon-graphite/ epoxy Hard Upper Torso (HUT)/Hatch assembly was designed, fabricated, and tested in the early 1990s. The spacesuit represented an 8.3 psi (≈58 kPa) technology demonstrator model of a zero prebreathe suit. The basic torso shell, brief, and hip areas of the suit were composed of a carbon-graphite/epoxy composite lay-up. In its current configuration, the suit weighs approximately 120 lb (≈54 kg). However, since future planetary suits will be designed to operate at 0.26 bar (≈26 kPa), it was felt that the suit's re-designed weight could be reduced to 79 lb (≈35 kg) with the incorporation of lightweight structural materials. Many robust, lightweight structures based on the technologies of advanced honeycomb materials, revolutionary new composite laminates, metal matrix composites, and recent breakthroughs in fullerene fillers and nanotechnology lend themselves well to applications requiring materials that are both light and strong. The major problem involves the reduction in weight of the HUT/ Hatch assembly for use in lunar and/or planetary applications, while at the same time maintaining a robust structural design. The technical objective is to research, design, and develop manufacturing methods that support fa b rica - tion of a lightweight HUT/Hatch assembly using advanced material and geometric redesign as necessary. Additionally, the lightweight HUT/Hatch assembly will interface directly with current MK III hardware. Using the new operating pressure and current MK III (H-1) interfaces as a starting block, it is planned to maximize HUT/Hatch assembly weight reduction through material selection and geometric redesign. A hard upper torso shell structure with rear-entry closure and corresponding hatch will be fabricated. The lightweight HUT/Hatch assembly will retrofit and interface with existing MK III (H-1) hardware elements, providing NASA with immediate "plug-andplay" capability. NASA crewmembers will have a lightweight, robust, life-support system that will minimize fatigue during extraterrestrial surface sojourns. Its unique feature is the utilization of a new and innovative family of materials used by the aerospace industry, which at the time of this reporting has not been used for the proposed application.

McCarthy, Mike↗

NASA NTRS1992

Scale effects on the transverse tensile strength of graphite epoxy composites

The influence of material volume on the transverse tensile strength of AS4/3501-6 graphite epoxy composites was investigated. Tensile tests of 90 degree laminates with 3 different widths and 5 different thicknesses were conducted. A finite element analysis was performed to determine the influence of the grip on the stress distribution in the coupons and explain the tendency for the distribution of failure locations to be skewed toward the grip. Specimens were instrumented with strain gages and extensometers to insure good alignment and to measure failure strains. Data indicated that matrix dominated strength properties varied with the volume of material that was stressed, with the strength decreasing as volume increased. Transverse strength data were used in a volumetric scaling law based on Weibull statistics to predict the strength of 90 degree laminates loaded in three point bending. Comparisons were also made between transverse strength measurements and out-of-plane interlaminar tensile strength measurements from curved beam bending tests. The significance of observed scale effects on the use of tests for material screening, quality assurance, and design allowables is discussed.

Obrien, T. Kevin↗

NASA NTRS1985

Response of thick, notched laminates subjected to tension-compression cyclic loads

The fatigue response of a (0/24/90/-45)s4 T300-5208 graphite-epoxy laminate with a drilled centerhole subjected to fully reversed tension-compression (R=-1, T-C) constant amplitude loading was investigated. Damage evaluation techniques such as stiffness monitoring, penetrant-enchanced X-ray radiography, C-scan, laminate deeply and residual strength were used to establish the mechanisms of damage development as well as the relations between this damage and the stiffness, strength and life of the laminate. Damage initiated at the hole as matrix cracking parallel to the fibers in all plies. Matrix cracks had a significant effect on delamination initiation and growth. Delaminations initiated near the surface in the densely cracked region at the hole and grew along major matrix cracks. Compressive properties degraded more rapidly than tensile properties. At the stress levels used, residual tensile strength increased early in the fatigue life and remained approximately constant to near the end of life, when failure was precipitated by excessive laminate instability during the compressive portion of the loading.

Bakis, C. E.↗

NASA NTRS1986

Response of thick, notched laminates subjected to tension-compression cyclic loads

The fatigue response of a (0/24/90/-45)s4 T300-5208 graphite-epoxy laminate with a drilled centerhole subjected to fully reversed tension-compression (R=-1, T-C) constant amplitude loading was investigated. Damage evaluation techniques such as stiffness monitoring, penetrant-enhanced X-ray radiography, C-scan, laminate deply and residual strength were used to establish the mechanisms of damage development as well as the relations between this damage and the stiffness, strength and life of the laminate. Damage initiated at the hole as matrix cracking parallel to the fibers in all plies. Matrix cracks had a significant effect on delamination initiation and growth. Delaminations initiated near the surface in the densely cracked region at the hole and grew along major matrix cracks. Compressive properties degraded more rapidly than tensile properties. At the stress levels used, residual tensile strength increased early in the fatigue life and remained approximately constant to near the end of life, when failure was precipitated by excessive laminate instability during the compressive portion of the loading.

Bakis, C. E.↗