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Results for “carbon-epoxy”

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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At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1983

Characterization of the matrix glass transition in carbon-epoxy laminates using the CSD test geometry

A new test geometry, referred to as centro-symmetric deformation (CSD), is proposed for characterizing the viscoelastic behavior of the matrix of carbon-epoxy laminates. The sample consists of a thin disk, typically 6-14 plies thick, having a nominal diameter of 30 mm. The disk is freely supported on a circular anvil; the load is applied to the center of the disk using an 8-mm-diameter ball bearing nosepiece. The CSD test geometry provides viscoelastic dispersion data which are independent of the angular orientation of the sample. The test geometry is sufficiently sensitive to matrix changes to allow its use for postcuring, humidity, crosslink density, and other matrix change studies. Test results are presented for a carbon-epoxy laminate.

Sternstein, S. S.↗
NASA NTRS1974

Comparison of fracture mechanisms for carbon-epoxy composites under tensile loading

The effects of oxidative surface treatments and differences in resin crack sensitivity on the tensile properties of carbon-epoxy composites have been investigated. A comparison of the fracture mechanisms for two different fiber batches was made in both model and engineering composites, using unmodified and modified epoxy resin. It has been observed that two distinct characteristics occur in the fracture surfaces of most engineering composites; first, fibers which fracture at the surface result in cleavage failure and, second, those which have fractured elsewhere and pulled out. A detailed discussion of these observations is contained in this paper.

Mazzio, V. F.↗
NASA NTRS1986

Lessons learned from the DC-10 carbon-epoxy rudder program

With the DC-10 carbon-epoxy rudder, the Douglas Aircraft Company achieved one of the greatest percentage weight savings associated with composite structures. Apart from minor damage from lightning strikes, the 15 rudders put into service have experienced virtually trouble-free operation for about a decade. This paper explains why a multirib, postbuckled skin design was used for the DC-10 composite rudder, how it was justified, and how it would have compared with more conventional sandwich design concepts. Special attention is devoted to the reasons why, for such postbuckled designs, it is better to allow the skin to wrinkle and unload itself than to reinforce it and make it resist buckling until some higher load level is attained. With minor changes in the manufacturing technique, this design concept is ideal for the control surfaces on many aircraft. The paper includes suggestions on how to make even better composite control surfaces in the future.

Hart-Smith, L. J.↗
NASA NTRS1989

Toughness properties of a three-dimensional carbon-epoxy composite

The three-dimensional (3D) orthogonal interlocked fabric contains through-the-thickness reinforcement in order to enhance the interlaminar fracture toughness of the composite. The interlaminar fracture toughness of a carbon-epoxy orthogonal interlocked fabric composite was experimentally determined by use of the recently developed tabbed double cantilever beam specimen. The data reduction methods applicable to these tests and materials and the interpretation of the results were discussed. The results of critical strain energy release rate, G(IC), were compared to those of a two-dimensional (2D) laminate having the same in-plane structure. The energy-dissipating crack propagation processes were described. The in-plane fracture toughness of the three-dimensional fabric was experimentally measured and compared to that of the two-dimensional laminate. The through-the-thickness fibers were found to create a tenfold increase in interlaminar toughness, and a 25-percent improvement in the in-plane fracture toughness.

Guenon, Valerie↗
NASA NTRS1971

Basic failure mechanisms in advanced composites

Fundamental failure mechanisms in carbon-epoxy composites were studied for more reliable prediction of the performance of these materials. Single and multiple fiber specimens were tested under tensile loads, and the sequence of failure events was observed. Parameters such as resin crack sensitivity, fiber surface treatment and variations in fibers from batch to batch are being evaluated. The analysis of bulk composite fracture processes using acoustic emission techniques is being studied in order to correlate microscopic observations with bulk composite behavior. Control of the fracture process through matrix and interface modification is being attempted, and study of failure processes in composite/metal specimens is being conducted. Most of the studies involved DEN 438 epoxy novolac as the matrix, but some experiments are now underway using the higher temperature resin ERLA 4617.

Mullin, J. V.↗
NASA NTRS1972

Basic failure mechanisms in advanced composites

Failure mechanisms in carbon-epoxy composites are identified as a basis for more reliable prediction of the performance of these materials. The approach involves both the study of local fracture events in model specimens containing small groups of filaments and fractographic examination of high fiber content engineering composites. Emphasis is placed on the correlation of model specimen observations with gross fracture modes. The effects of fiber surface treatment, resin modification and fiber content are studied and acoustic emission methods are applied. Some effort is devoted to analysis of the failure process in composite/metal specimens.

Mullin, J. V.↗
NASA NTRS1973

Basic failure mechanisms in advanced composites

The fundamental failure mechanisms which result from the interaction of thermal cycling and mechanical loading of carbon-epoxy composites were studied. This work was confined to epoxy resin uniderictionally reinforced with HTS carbon fibers, and consists of first identifying local fiber, matrix and interface failure mechanisms using the model composite specimen containing a small number of fibers so that optical techniques can be used for characterization. After the local fracture process has been established for both mechanical loading and thermal cycling, engineering composite properties and gross fracture modes are then examined to determine how the local events contribute to real composite performance. Flexural strength in high fiber content specimens shows an increase in strength with increased thermal cycling. Similar behavior is noted for 25 v/o material up to 200 cycles; however, there is a drastic reduction after 200 cycles indicating a major loss of integrity probably through the accumulation of local cleavage cracks in the tensile region.

Mazzio, V. F.↗
NASA NTRS1985

The effect of Low Earth Orbit space environment on polymeric spacecraft materials

The Advanced Composite Materials Exposure to Space Experiment flown on shuttle mission STS-41G was designed to investigate the effect of low Earth orbit space environment on polymeric spacecraft materials. The effect of the phenomenon attributed to atomic oxygen on Kapton and advanced composite materials of carbon-epoxy and kevlar epoxy is described. The results of true space exposure are compared to duplicate specimens subjected to a simulated environment in an Earth-based nozzle beam test facility utilizing microwave discharge to form atomic oxygen. Although the energy level for atomic oxygen in the beam facility did not achieve the orbital energy level of 5 eV, similar effects of surface erosion and degradation are observed after equivalent exposure times.

Zimcik, D. G.↗
NASA NTRS1985

Results of apparent atomic oxygen reactions with spacecraft materials during shuttle flight STS-41G

The effect of atomic oxygen interaction experienced by polymeric-based spacecraft materials is described. An experimental package (ACOMEX) flown on shuttle mission STS-41G carried out the investigation of advanced composite specimens such as carbon-epoxy and Kevlar-epoxy both with and without protective coatings added to thermal protective paints and films. Information on the exposure environment of the specimens was provided by a carbon coated atomic oxygen fluence monitor together with a photographic record. Mass loss measurements and photomicrographs made possible the analysis of the effect of interaction. After a total of about 38 hours of equivalent normal exposure at 225 km altitude the results showed that unprotected exposed surfaces exhibited severe erosion and mass loss with the possibility of seriously degrading structural and thermal performance. However, the specimens with a thin fluorocarbon overcoat showed promise of providing a protective barrier to the attack without altering the base properties of the material.

Zimcik, D. G.↗
NASA NTRS1986

Ten years of flight service with DC-10 composite rudders - A backward glance

Carbon-epoxy upper aft rudders have been successfully evaluated on the DC-10 during 10 years of airline flight service. They have flown for more than 320,000 flight hours, with 36,000 flight hours flown by a single rudder. The rudder box is of thin-skin buckled construction and is cocured out-of-autoclave as a one-piece assembly. Weight saving can exceed the rudder weight if full advantage is taken of removing balance weight from the forward rudder.

Hawley, Arthur V.↗
NASA NTRS1988

The Purdue University Get Away Special II (PUGAS II)

The Purdue University Get Away Special Project (PUGAS) is a student-run organization dedicated to preparing payloads for flight on NASA's space shuttle. The first such payload (PUGAS I) flew on Challenger in 1983. The second payload (PUGAS II) should be ready by the end of 1988 and will include three experiments. The first experiment will involve the production of tin metal foam under microgravity conditions. The second experiment will focus on the desorption of water from carbon-epoxy composite materials. The third experiment will use a solid polymeric material to detect radiation in space.

Olenski, Christopher↗
NASA NTRS1989

The soft x ray telescope for Solar-A

The Solar-A satellite being prepared by the Institute for Sapce and Astronautical Sciences (ISAS) in Japan is dedicated to high energy observations of solar flares. The Soft X Ray Telescope (SXT) is being prepared to provide filtered images in the 2 to 60 A interval. The flight model is now undergoing tests in the 1000 foot tunnel at MSFC. Launch will be in September 1991. Earlier resolution and efficiency tests on the grazing incidence mirror have established its performance in soft x rays. The one-piece, two mirror grazing incidence telescope is supported in a strain free mount separated from the focal plane assembly by a carbon-epoxy metering tube whose windings and filler are chosen to minimize thermal and hygroscopic effects. The CCD detector images both the x ray and the concentric visible light aspect telescope. Optical filters provide images at 4308 and 4700 A. The SXT will be capable of producing over 8000 of the smallest partial frame images per day, or fewer but larger images, up to 1024 x 1024 pixel images. Image sequence with two or more of the five x ray analysis filters, with automatic exposure compensation to optimize the charge collection by the CCD detector, will be used to provide plasma diagnostics. Calculations using a differential emission measure code were used to optimize filter selection over the range of emission measure variations and to avoid redundancy, but the filters were chosen primarily to give ratios that are monotonic in plasma temperature.

Brown, W. A.↗
NASA NTRS1990

Instrumented impact and residual tensile strength testing of eight-ply carbon eopoxy specimens

Instrumented drop weight impact testing was utilized to examine a puncture-type impact on thin carbon-epoxy coupons. Four different material systems with various eight-ply lay-up configurations were tested. Specimens were placed over a 10.3-mm diameter hole and impacted with a smaller tup (4.2-mm diameter) than those used in previous studies. Force-time plots as well as data on absorbed energy and residual tensile strength were gathered and examined. It was found that a critical impact energy level existed for each material tested, at which point tensile strength began to rapidly decrease with increasing impact energy.

Nettles, A. T.↗
NASA NTRS1990

An examination of impact damage in glass-phenolic and aluminum honeycomb core composite panels

An examination of low velocity impact damage to glass-phenolic and aluminum core honeycomb sandwich panels with carbon-epoxy facesheets is presented. An instrumented drop weight impact test apparatus was utilized to inflict damage at energy ranges between 0.7 and 4.2 joules. Specimens were checked for extent of damage by cross sectional examination. The effect of core damage was assessed by subjecting impact-damaged beams to four-point bend tests. Skin-only specimens (facings not bonded to honeycomb) were also tested for comparison purposes. Results show that core buckling is the first damage mode, followed by delaminations in the facings, matrix cracking, and finally fiber breakage. The aluminum honeycomb panels exhibited a larger core damage zone and more facing delaminations than the glass-phenolic core, but could withstand more shear stress when damaged than the glass-phenolic core specimens.

Nettles, A. T.↗
NASA NTRS1990

Mode I delamination of a three-dimensional fabric composite

Finite-element analysis was conducted on a double-cantilever-beam (DCB) specimen of carbon-epoxy (T300/3501-6) orthogonal interlocked fabric composites. The mode I strain-energy release rate G(I) was evaluated to investigate the influence of through-the-thickness (z-axis) fibers on crack driving force as a function of crack length. The analysis showed good agreement with experimental results for both compliance and G(I). The presence of z-axis fibers sharply reduced G(I) in comparison to the traditional two-dimensional laminated material. A parametric study revealed that increasing the z-axis fiber stiffness reduces G(I). Progressive debonding of the z-axis reinforcement was also modeled. G(I) increased as the fiber debonded but was significantly less than in the two-dimensional laminated material in all cases.

Byun, Joon-Hyung↗
NASA NTRS1991

Materials and processes laboratory composite materials characterization task, part 1. Damage tolerance

A test run was performed on IM6/3501-6 carbon-epoxy in which the material was processed, machined into specimens, and tested for damage tolerance capabilities. Nondestructive test data played a major role in this element of composite characterization. A time chart was produced showing the time the composite material spent within each Branch or Division in order to identify those areas which produce a long turnaround time. Instrumented drop weight testing was performed on the specimens with nondestructive evaluation being performed before and after the impacts. Destructive testing in the form of cross-sectional photomicrography and compression-after-impact testing were used. Results show that the processing and machining steps needed to be performed more rapidly if data on composite material is to be collected within a reasonable timeframe. The results of the damage tolerance testing showed that IM6/3501-6 is a brittle material that is very susceptible to impact damage.

Nettles, A. T.↗