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  • 1
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    Effects of transverse shearing on cylindrical bending, vibration, and buckling of laminated plates (1987)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Keywords: STRUCTURAL MECHANICS
    Type: AIAA Journal (ISSN 0001-1452); 25; 123-129
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    A study of graphite-epoxy laminate failures due to high transverse shear strains using the multi-span-beam shear test procedure (1989)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: The multi-span-beam shear test procedure is used to study failure mechanisms in graphite-epoxy laminates due to high transverse shear strains induced by severe local bending deformations in test specimens. Results of a series of tests on specimens with a variety of stacking sequences, including some with adhesive interleaving, are presented. These results indicate that laminates with stacking sequences with several + or - 45 and 90 deg plies next to each other are more susceptible to failures due to high transverse shear strains than laminates with + or - 45 and 0 deg plies next to each other or with + or - 45 deg plies next to layers of adhesive interleaving. Results of these tests are compared with analytical results based on finite elements.
    Keywords: COMPOSITE MATERIALS
    Format: text
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    NASA TECHNICAL REPORTS
  • 3
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    Study of thermal-expansion-molded, graphite-epoxy hat-stiffened sandwich panels (1991)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Keywords: STRUCTURAL MECHANICS
    Type: AIAA Journal (ISSN 0001-1452); 29; 778
    Format: text
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    NASA TECHNICAL REPORTS
  • 4
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    Effect of composite fabrication method on structural response and impact damage (1992)
    In:  Other Sources
    Details
    Publication Date: 2011-08-24
    Keywords: COMPOSITE MATERIALS
    Type: AIAA Journal (ISSN 0001-1452); 30; 205-213
    Format: text
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    NASA TECHNICAL REPORTS
  • 5
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    Comparison of hand laid-up tape and filament wound composite cylinders and panels with and without impact damage (1991)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The results of this experimental comparison of filament wound control (unimpacted) cylinders loaded to failure in axial compression indicates that one fiber cross-over location has no effect on the failure mode or strain in thick walled filament wound graphite-epoxy specimens with stacking sequence (plus or minus 45/90) sub 3s. A comparison between filament wound and hand laid-up tape control cylinders indicates that there is little or no difference in the response of cylinders constructed by using two different fabrication methods, however, unimpacted panels with many fiber cross-overs fail at up to 15 percent lower strains than panels with no fiber cross-overs. A comparison of samples subjected to low speed impact damage prior to compressive loading indicates that impact damage reduces the strain at failure by over 60 percent in tape and filament wound graphite-epoxy cylinders and in tape flat panels. The presence of fiber cross-overs was observed to reduce the strength of filament wound impact-damaged panels, but to have no significant effect on the strength of filament wound impact-damaged cylinders.
    Keywords: COMPOSITE MATERIALS
    Type: First NASA Advanced Composites Technology Conference, Part 2; p 749-759
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    A study of structurally efficient graphite-thermoplastic trapezoidal-corrugation sandwich and semi-sandwich panels (1993)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The structural efficiency of compression-loaded trapezoidal-corrugation sandwich and semi-sandwich composite panels is studied to determine their weight savings potential. Sandwich panels with two identical face sheets and a trapezoidal corrugated core between them, and semi-sandwich panels with a corrugation attached to a single skin are considered. An optimization code is used to find the minimum weight designs for critical compressive load levels ranging from 3,000 to 24,000 lb/in. Graphite-thermoplastic panels based on the optimal minimum weight designs were fabricated and tested. A finite-element analysis of several test specimens was also conducted. The results of the optimization study, the finite-element analysis, and the experiments are presented.
    Keywords: COMPOSITE MATERIALS
    Type: Third NASA Advanced Composites Technology Conference, Volume 1, Part 2; p 859-878
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  • 7
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    Compression behavior of graphite-epoxy and graphite-thermoplastic panels with circular holes or impact damage (1990)
    In:  CASI
    Details
    Publication Date: 2013-08-29
    Description: An experimental investigation of the compression behavior of laminated specimens made from graphite-epoxy tape (AS4-3502), graphite-thermoplastic tape (AS4-PEEK), and graphite-thermoplastic fabric (AS4-PEEK) 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 energy up to 30 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 of fabric indicate that the material form has little effect on failure strains in specimens with holes or low speed impact damage.
    Keywords: COMPOSITE MATERIALS
    Type: Eighth DOD(NASA)FAA Conference on Fibrous Composites in Structural Design, Part 2; p 537-558
    Format: text
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  • 8
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    Structural Test Documentation and Results for the McDonnell Douglas All-Composite Wing Stub Box (1997)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The results of a series of tests conducted at the NASA Langley Research Center to evaluate the behavior of an all-composite full-scale wing box are presented. The wing stub box is representative of a section of a commercial transport aircraft wing box and was designed and constructed by McDonnell Douglas Aerospace Company as part of the NASA Advanced Composites Technology (ACT) program. Tests were conducted with and without low-speed impact damage and repairs. The structure with nonvisible impact damage carried 140 percent of Design Limit Load prior to failure through an impact site.
    Keywords: Structural Mechanics
    Type: NASA-TM-110204 , NAS 1.15:110204
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 9
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    Correlation of Structural Analysis and Test Results for the McDonnell Douglas Stitched/RFI All-Composite Wing Stub Box (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The analytical and experimental results of an all-composite wing stub box are presented in this report. The wing stub box, which is representative of an inboard portion of a commercial transport high-aspect-ratio wing, was fabricated from stitched graphite-epoxy material with a Resin Film Infusion manufacturing process. The wing stub box was designed and constructed by the McDonnell Douglas Aerospace Company as part of the NASA Advanced Composites Technology program. The test article contained metallic load-introduction structures on the inboard and outboard ends of the graphite-epoxy wing stub box. The root end of the inboard load introduction structure was attached to a vertical reaction structure, and an upward load was applied to the outermost tip of the outboard load introduction structure to induce bending of the wing stub box. A finite element model was created in which the center portion of the wing-stub-box upper cover panel was modeled with a refined mesh. The refined mesh was required to represent properly the geometrically nonlinear structural behavior of the upper cover panel and to predict accurately the strains in the stringer webs of the stiffened upper cover panel. The analytical and experimental results for deflections and strains are in good agreement.
    Keywords: Structural Mechanics
    Type: NASA-TM-110267 , NAS 1.15:110267
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 10
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    Analysis of selected compression splice joint locations in a graphite-epoxy transport wing stub box (1995)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Three critical compression splice joint locations in a stitched graphite-epoxy transport wing stub box have been analyzed to determine their expected structural performance. The wing box is representative of a section of a commercial transport wing box and was designed and constructed by McDonnell Douglas Aerospace Company as part of the NASA Advanced Composites Technology (ACT) program. The results of the finite element analyses of the splice joints are presented. The analysis results indicate that failure will not occur in the splice joint regions for loads less than the Design Ultimate Load of the wing box.
    Keywords: STRUCTURAL MECHANICS
    Type: NASA-TM-110170 , NAS 1.15:110170 , NIPS-95-06489
    Format: application/pdf
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