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  • Other Sources  (82)
  • COMPOSITE MATERIALS
  • 1995-1999  (82)
  • 1990-1994
  • 1995  (82)
  • 1
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    Effects of Thermal Treatment on Tensile Creep and Stress-Rupture Behavior of Hi-Nicalon SiC Fibers (1995)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Tensile creep and stress-rupture studies were conducted on Hi-Nicalon SiC fibers at 1200 and 1400 C in argon and air. Examined were as-received fibers as well as fibers annealed from 1400 to 1800 C for 1 hour in argon before testing. The creep and rupture results for these annealed fibers were compared to those of the as-received fibers to determine the effects of annealing temperature, test temperature, and test environment. Argon anneals up to 1500 C degrade room temperature strength of Hi-Nicalon fibers, but improve fiber creep resistance in argon or air by as much as 100% with no significant degradation in rupture strength. Argon anneals above 1500 C continue to improve fiber creep resistance when tested in argon, but significantly degrade creep resistance and rupture strength when tested in air. Decrease in creep resistance in air is greater at 1200 C than at 1400 C. Mechanisms are suggested for the observed behavior.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-107087 , NAS 1.15:107087 , E-9965 , Annual Conference on Composites, Advanced Ceramics, Materials and Structures; Jan 08, 1995 - Jan 15, 1995; Cocoa Beach, FL; United States
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  • 2
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    Micromechanics of composites with shape memory alloy fibers in uniform thermal fields (1995)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Analytical procedures are developed for a composite system consisting of shape memory alloy fibers within an elastic matrix subject to uniform temperature fluctuations. Micromechanics for the calculation of the equivalent properties of the composite are presented by extending the multi-cell model to incorporate shape memory alloy fibers. A three phase concentric cylinder model is developed for the analysis of local stresses which includes the fiber, the matrix, and the surrounding homogenized composite. The solution addresses the complexities induced by the nonlinear dependence of the in-situ martensite fraction of the fibers to the local stresses and temperature, and the local stresses developed from interactions between the fibers and matrix during the martensitic and reverse phase transformations. Results are presented for a nitinol/epoxy composite. The applications illustrate the response of the composite in isothermal longitudinal loading and unloading, and in temperature induced actuation. The local stresses developed in the composite under various stages of the martensitic and reverse phase transformation are also shown.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-107011 , E-9795 , NAS 1.15:107011 , AIAA PAPER 95-1210 , Structures, Structural Dynamics and Materials Conference; Apr 10, 1995 - Apr 13, 1995; New Orleans, LA; United States
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  • 3
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    Formulation of an improved smeared stiffener theory for buckling analysis of grid-stiffened composite panels (1995)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A smeared stiffener theory for stiffened panels is presented that includes skin-stiffener interaction effects. The neutral surface profile of the skin-stiffener combination is developed analytically using the minimum potential energy principle and statics conditions. The skin-stiffener interaction is accounted for by computing the stiffness due to the stiffener and the skin in the skin-stiffener region about the neutral axis at the stiffener. Buckling load results for axially stiffened, orthogrid, and general grid-stiffened panels are obtained using the smeared stiffness combined with a Rayleigh-Ritz method and are compared with results from detailed finite element analyses.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-110162 , NAS 1.15:110162 , International Conference on Composite Materials; Aug 14, 1995 - Aug 18, 1995; British Columbia; Canada
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  • 4
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    EMI Shields made from intercalated graphite composites (1995)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Electromagnetic interference (EMI) shielding typically makes up about twenty percent of the mass of a spacecraft power system. Graphite fiber/polymer composites have significantly lower densities and higher strengths than aluminum, the present material of choice for EMI shields, but they lack the electrical conductivity that enables acceptable shielding effectiveness. Bromine intercalated pitch-based graphite/epoxy composites have conductivities fifty times higher than conventional structural graphite fibers. Calculations are presented which indicate that EMI shields made from such composites can have sufficient shielding at less than 20% of the mass of conventional aluminum shields. EMI shields provide many functions other than EMI shielding including physical protection, thermal management, and shielding from ionizing radiation. Intercalated graphite composites perform well in these areas also. Mechanically, they have much higher specific strength and modulus than aluminum. They also have shorter half thicknesses for x-rays and gamma radiation than aluminum. Thermally, they distribute infra-red radiation by absorbing and re-radiating it rather than concentrating it by reflection as aluminum does. The prospects for intercalated graphite fiber/polymer composites for EMI shielding are encouraging.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-106979 , E-9743 , NAS 1.15:106979 , International Materials and Processing Conference; Jun 21, 1994 - Jun 23, 1994; Parsippany, NJ; United States
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  • 5
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    Design and evaluation of a foam-filled hat-stiffened panel concept for aircraft primary structural applications (1995)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A structurally efficient hat-stiffened panel concept that utilizes a structural foam as stiffener core has been designed for aircraft primary structural applications. This stiffener concept utilizes a manufacturing process that can be adapted readily to grid-stiffened structural configurations which possess inherent damage tolerance characteristics due to their multiplicity of load paths. The foam-filled hat-stiffener concept in a prismatically stiffened panel configuration is more efficient than most other stiffened panel configurations in a load range that is typical for both fuselage and wing structures. The prismatically stiffened panel concept investigated here has been designed using AS4/3502 preimpregnated tape and Rohacell foam core and evaluated for its buckling and postbuckling behavior with and without low-speed impact damage. The results from single-stiffener and multi-stiffener specimens suggest that this structural concept responds to loading as anticipated and has good damage tolerance characteristics.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-109175 , NAS 1.15:109175 , Advanced Composites Technology Conference; Jun 01, 1992
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  • 6
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    The effects of tensile preloads on the impact response of carbon/epoxy laminates (1995)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The effects of tensile preloads on the tension-after-impact (TAI) strength of composite laminates of IM7/8551-7 were examined. A failure threshold curve was first determined so the most informative values for preload/impact energy combinations could be determined. The impact tests were instrumented so maximum load of impact, as well as several other parameters could be measured. The elastic response data indicate that as the tensile preload is increased, the maximum load of impact also increases. The damage data show that at low impact energies, the damage/failure is an 'all-or-nothing' event but at higher impact energies, a region of preload values exists where the coupons could sustain damage, yet not fail catastrophically.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-111356 , NAS 1.15:111356 , NIPS-96-08313 , International SAMPE Technical Conference; May 08, 1995 - May 11, 1995; Anaheim, CA; United States
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  • 7
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    The Effects of Fiber Surface Modification and Thermal Aging on Composite Toughness and Its Measurement (1995)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A detailed experimental study was conducted to establish the structure-property relationships between elevated temperature aging and fiber-matrix bonding, Mode 2 interlaminar fracture toughness, and failure modes of carbon fiber/PMR-15 composites. The fiber-matrix adhesion was varied by using carbon fibers with different surface treatments. Short beam shear tests were used to quantify the interfacial shear strength afforded by the use of the different fiber surface treatments. The results of the short beam shear tests showed that, for times up to 1000 hr, the aging process caused no changes in the bulk of the three composite materials that would degrade the shear properties of the material. Comparisons between the interlaminar shear strengths (ILSS) measured by the short beam shear tests and the GIIC test results, as measured by the ENF test, indicated that the differences in the surface treatments significantly affected the fracture properties while the effect of the aging process was probably limited to changes at the starter crack tip. The fracture properties changed due to a shift in the fracture from an interfacial failure to a failure within the matrix when the fiber was changed from AU-4 to AS-4 or AS-4G. There appears to be an effect of the fiber/matrix bonding on the thermo-oxidative stability of the composites that were tested. The low bonding afforded by the AU 1 fiber resulted in weight losses about twice those experienced by the AS 1 reinforced composites, the ones with the best TOS.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-106765 , NAS 1.15:106765 , E-9816 , NIPS-96-08132 , International Symposium and Exhibition; May 08, 1995 - May 11, 1995; Anaheim, CA; United States
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  • 8
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    Verification of a three-dimensional resin transfer molding process simulation model (1995)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Experimental evidence was obtained to complete the verification of the parameters needed for input to a three-dimensional finite element model simulating the resin flow and cure through an orthotropic fabric preform. The material characterizations completed include resin kinetics and viscosity models, as well as preform permeability and compaction models. The steady-state and advancing front permeability measurement methods are compared. The results indicate that both methods yield similar permeabilities for a plain weave, bi-axial fiberglass fabric. Also, a method to determine principal directions and permeabilities is discussed and results are shown for a multi-axial warp knit preform. The flow of resin through a blade-stiffened preform was modeled and experiments were completed to verify the results. The predicted inlet pressure was approximately 65% of the measured value. A parametric study was performed to explain differences in measured and predicted flow front advancement and inlet pressures. Furthermore, PR-500 epoxy resin/IM7 8HS carbon fabric flat panels were fabricated by the Resin Transfer Molding process. Tests were completed utilizing both perimeter injection and center-port injection as resin inlet boundary conditions. The mold was instrumented with FDEMS sensors, pressure transducers, and thermocouples to monitor the process conditions. Results include a comparison of predicted and measured inlet pressures and flow front position. For the perimeter injection case, the measured inlet pressure and flow front results compared well to the predicted results. The results of the center-port injection case showed that the predicted inlet pressure was approximately 50% of the measured inlet pressure. Also, measured flow front position data did not agree well with the predicted results. Possible reasons for error include fiber deformation at the resin inlet and a lag in FDEMS sensor wet-out due to low mold pressures.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-CR-199792 , NAS 1.26:199792 , CCMS-95-10 , VPI-E-95-04 , NIPS-95-06462
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  • 9
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    Optimization of composite box-beam structures including effects of subcomponent interactions (1995)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Minimum mass designs are obtained for a simple box beam structure subject to bending, torque and combined bending/torque load cases. These designs are obtained subject to point strain and linear buckling constraints. The present work differs from previous efforts in that special attention is payed to including the effects of subcomponent panel interaction in the optimal design process. Two different approaches are used to impose the buckling constraints. When the global approach is used, buckling constraints are imposed on the global structure via a linear eigenvalue analysis. This approach allows the subcomponent panels to interact in a realistic manner. The results obtained using this approach are compared to results obtained using a traditional, less expensive approach, called the local approach. When the local approach is used, in-plane loads are extracted from the global model and used to impose buckling constraints on each subcomponent panel individually. In the global cases, it is found that there can be significant interaction between skin, spar, and rib design variables. This coupling is weak or nonexistent in the local designs. It is determined that weight savings of up to 7% may be obtained by using the global approach instead of the local approach to design these structures. Several of the designs obtained using the linear buckling analysis are subjected to a geometrically nonlinear analysis. For the designs which were subjected to bending loads, the innermost rib panel begins to collapse at less than half the intended design load and in a mode different from that predicted by linear analysis. The discrepancy between the predicted linear and nonlinear responses is attributed to the effects of the nonlinear rib crushing load, and the parameter which controls this rib collapse failure mode is shown to be the rib thickness. The rib collapse failure mode may be avoided by increasing the rib thickness above the value obtained from the (linear analysis based) optimizer. It is concluded that it would be necessary to include geometric nonlinearities in the design optimization process if the true optimum in this case were to be found.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-CR-199667 , NAS 1.26:199667 , CCMS-95-06 , VPI-E-95-03 , NIPS-95-05666
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  • 10
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    Materials research for high-speed civil transport and generic hypersonics: Composites durability (1995)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This report covers a portion of an ongoing investigation of the durability of composites for the High Speed Civil Transport (HSCT) program. Candidate HSCT composites need to possess the high-temperature capability required for supersonic flight. This program was designed to initiate the design, analysis, fabrication, and testing of equipment intended for use in validating the long-term durability of materials for the HSCT. This equipment includes thermally actuated compression and tension fixtures, hydraulic-actuated reversible load fixtures, and thermal chambers. This equipment can be used for the durability evaluation of both composite and adhesive materials. Thermally actuated fixtures are recommended for fatigue cycling when long-term thermomechanical fatigue (TMF) data are required on coupon-sized tension or compression specimens. Long term durability testing plans for polymer matrix composite specimens are included.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-CR-198192 , NAS 1.26:198192
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