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  • 1
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    Nonlinear substructuring method for concurrent processing computers (1987)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Keywords: STRUCTURAL MECHANICS
    Type: AIAA Journal (ISSN 0001-1452); 25; 871-876
    Format: text
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  • 2
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    A nonlinear substructuring method for concurrent processing computers (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: This paper proposes a method, based on substructuring, to solve nonlinear structural analysis problems on multiprocessor computers. Background information is given on the use of substructuring in large-scale finite element programs and computational time distributions for the major components for an example nonlinear finite element analysis are discussed. Implementation of the substructuring method on a typical multiprocessor computer is described and estimates are made of expected reductions in computation times based on nonlinear substructuring results obtained on a single processor computer.
    Keywords: STRUCTURAL MECHANICS
    Type: AIAA PAPER 86-0852
    Format: text
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  • 3
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    Development of a Mesoscale Finite Element Constitutive Model for Fiber Kinking (2018)
    In:  CASI
    Details
    Publication Date: 2019-07-20
    Description: A mesoscale finite element material model is proposed to analyze structures that fail by the fiber kinking damage mode. To evaluate the assumptions of the mesoscale model, the results were compared with those of a high-fidelity micromechanical model. A direct comparison between the two models shows remarkable correlation, indicating that the key features of the fiber kinking phenomenon are appropriately accounted for in the mesoscale model. The mesoscale model is applied to structural analysis cases to demonstrate the capabilities of the model. A verification study is conducted with an unnotched compression specimen and preliminary validation is demonstrated with a notched compression specimen. The results show that the model is successful at representing the kinematics of fiber kinking while at the same time highlighting the need for further verification and validation.
    Keywords: Composite Materials
    Type: NF1676L-27418 , AIAA SciTech 2018; Jan 08, 2018 - Jan 12, 2018; Kissimmee, FL; United States
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  • 4
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    Mode I Cohesive Law Characterization of Through-Crack Propagation in a Multidirectional Laminate (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A method is proposed and assessed for the experimental characterization of through-the-thickness crack propagation in multidirectional composite laminates with a cohesive law. The fracture toughness and crack opening displacement are measured and used to determine a cohesive law. Two methods of computing fracture toughness are assessed and compared. While previously proposed cohesive characterizations based on the R-curve exhibit size effects, the proposed approach results in a cohesive law that is a material property. The compact tension specimen configuration is used to propagate damage while load and full-field displacements are recorded. These measurements are used to compute the fracture toughness and crack opening displacement from which the cohesive law is characterized. The experimental results show that a steady-state fracture toughness is not reached. However, the proposed method extrapolates to steady-state and is demonstrated capable of predicting the structural behavior of geometrically-scaled specimens.
    Keywords: Composite Materials
    Type: NF1676L-18222 , US-Japan Conference on Composite Materials; Sep 08, 2014 - Sep 10, 2014; La Jolla, CA; United States|ASTM-D30 Meeting; Sep 08, 2014 - Sep 10, 2014; La Jolla, CA; United States|American Society for Composites Technical Conference; Sep 08, 2014 - Sep 10, 2014; La Jolla, CA; United States
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  • 5
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    Nonlinear Analysis and Post-Test Correlation for a Curved PRSEUS Panel (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept, developed by The Boeing Company, has been extensively studied as part of the National Aeronautics and Space Administration's (NASA s) Environmentally Responsible Aviation (ERA) Program. The PRSEUS concept provides a light-weight alternative to aluminum or traditional composite design concepts and is applicable to traditional-shaped fuselage barrels and wings, as well as advanced configurations such as a hybrid wing body or truss braced wings. Therefore, NASA, the Federal Aviation Administration (FAA) and The Boeing Company partnered in an effort to assess the performance and damage arrestments capabilities of a PRSEUS concept panel using a full-scale curved panel in the FAA Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) facility. Testing was conducted in the FASTER facility by subjecting the panel to axial tension loads applied to the ends of the panel, internal pressure, and combined axial tension and internal pressure loadings. Additionally, reactive hoop loads were applied to the skin and frames of the panel along its edges. The panel successfully supported the required design loads in the pristine condition and with a severed stiffener. The panel also demonstrated that the PRSEUS concept could arrest the progression of damage including crack arrestment and crack turning. This paper presents the nonlinear post-test analysis and correlation with test results for the curved PRSEUS panel. It is shown that nonlinear analysis can accurately calculate the behavior of a PRSEUS panel under tension, pressure and combined loading conditions.
    Keywords: Aircraft Design, Testing and Performance
    Type: AIAA Paper 2013-1736 , NF1676L-15294 , 54th AIAA/ASME/ASCE/AHS/ASC, Structures, Structural Dynamics, and Materials Conference; Apr 08, 2013 - Apr 11, 2013; Boston, MA; United States
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  • 6
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    Full-Scale Test and Analysis Results of a PRSEUS Fuselage Panel to Assess Damage Containment Features (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Integrally stitched composite technology is an area that shows promise in enhancing the structural integrity of aircraft and aerospace structures. The most recent generation of this technology is the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept. The goal of the PRSEUS concept relevant to this test is to provide damage containment capability for composite structures while reducing overall structural weight. The National Aeronautics and Space Administration (NASA), the Federal Aviation Administration (FAA), and The Boeing Company have partnered in an effort to assess the damage containment features of a full-scale curved PRSEUS panel using the FAA Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) facility. A single PRSEUS test panel was subjected to axial tension, internal pressure, and combined axial tension and internal pressure loads. The test results showed excellent performance of the PRSEUS concept. No growth of Barely Visible Impact Damage (BVID) was observed after ultimate loads were applied. With a two-bay notch severing the central stringer, damage was contained within the two-bay region well above the required limit load conditions. Catastrophic failure was well above the ultimate load level. Information describing the test panel and procedure has been previously presented, so this paper focuses on the experimental procedure, test results, nondestructive inspection results, and preliminary test and analysis correlation.
    Keywords: Aircraft Design, Testing and Performance
    Type: NF1676L-14599 , 2012 Aircraft Airworthiness and Sustainment Conference; Apr 02, 2012 - Apr 05, 2012; Baltimore, MD; United States
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  • 7
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    Assessment of Damage Containment Features of a Full-Scale PRSEUS Fuselage Panel Through Test and Teardown (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An area that shows promise in enhancing structural integrity of aircraft and aerospace structures is the integrally stitched composite technology. The most recent generation of this technology is the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept developed by Boeing Research and Technology and the National Aeronautics and Space Administration. A joint test program on the assessment of damage containment capabilities of the PRSEUS concept for curved fuselage structures was conducted recently at the Federal Aviation Administration William J. Hughes Technical Center. The panel was subjected to axial tension, internal pressure, and combined axial tension and internal pressure load conditions up to fracture, with a through-the-thickness, two-bay notch severing the central stiffener. For the purpose of future progressive failure analysis development and verification, extensive post failure nondestructive and teardown inspections were conducted. Detailed inspections were performed directly ahead of the notch tip where stable damage progression was observed. These examinations showed: 1) extensive delaminations developed ahead of the notch tip, 2) the extent and location of damage, 3) the typical damage mechanisms observed in composites, and 4) the role of stitching and warp-knitting in the failure mechanisms. The objective of this paper is to provide a summary of results from these posttest inspections.
    Keywords: Aircraft Design, Testing and Performance
    Type: NF1676L-14533 , 2012 American Society for Composites 27th Technical Conference; Oct 01, 2012 - Oct 03, 2012; Arlington, TX; United States|15th US-Japan Conference on Composite Materials; Oct 01, 2012 - Oct 03, 2012; Arlington, TX; United States
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  • 8
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    Full-Scale Test and Analysis of a PRSEUS Fuselage Panel to Assess Damage-Containment Features (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Stitched composite technology has the potential to substantially decrease structural weight through enhanced damage containment capabilities. The most recent generation of stitched composite technology, the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept, has been shown to successfully arrest damage at the sub-component level through tension testing of a three stringer panel with damage in the form of a two-bay notch. In a joint effort undertaken by the National Aeronautics and Space Administration (NASA), the Federal Aviation Administration (FAA), and the Boeing Company, further studies are being conducted to characterize the damage containment features of the PRSEUS concept. A full-scale residual strength test will be performed on a fuselage panel to determine if the load capacity will meet strength, deformation, and damage tolerance requirements. A curved panel was designed, fabricated, and prepared for residual strength testing. A pre-test Finite Element Model (FEM) was developed using design allowables from previous test programs to predict test panel deformation characteristics and margins of safety. Three phases of testing with increasing damage severity include: (1) as manufactured; (2) barely visible impact damage (BVID) and visible impact damage (VID); and (3) discrete source damage (DSD) where the panel will be loaded to catastrophic failure. This paper presents the background information, test plan, and experimental procedure. This paper is the first of several future articles reporting the test preparations, results, and analysis conducted in the test program.
    Keywords: Aircraft Design, Testing and Performance
    Type: NF1676L-11521 , 2011 Aircraft Airworthiness and Sustainment (AAS) Conference; Apr 18, 2011 - Apr 21, 2011; San Diego, CA; United States
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  • 9
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    Computational Performance of Progressive Damage Analysis of Composite Laminates using Abaqus/Explicit with 16 to 512 CPU Cores (2019)
    In:  CASI
    Details
    Publication Date: 2019-07-20
    Description: The computational scaling performance of progressive damage analysis using Abaqus/ Explicit is evaluated and quantified using from 16 to 512 CPU cores. Several analyses were conducted on varying numbers of cores to determine the scalability of the code on five NASA high performance computing systems. Two finite element models representative of typical models used for progressive damage analysis of composite laminates were used. The results indicate a 10 to 15 times speed up scaling from 24 to 512 cores. The run times were modestly reduced with newer generations of CPU hardware. If the number of degrees of freedom is held constant with respect to the number of cores, the model size can be increased by a factor of 20, scaling from 16 to 512 cores, with the same run time. An empirical expression was derived relating run time, the number of cores, and the number of degrees of freedom. Analysis cost was examined in terms of software tokens and hardware utilization. Using additional cores reduces token usage since the computational performance increases more rapidly than the token requirement with increasing number of cores. The in- crease in hardware cost with increasing cores was found to be modest. Overall the results show relatively good scalability of the Abaqus/Explicit code on up to 512 cores.
    Keywords: Computer Programming and Software; Composite Materials
    Type: NASA/TM-2019-220251 , L-20998 , NF1676L-32380
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  • 10
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    In-Situ Observations of Longitudinal Compression Damage in Carbon-Epoxy Cross Ply Laminates Using Fast Synchrotron Radiation Computed Tomography (2017)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The role of longitudinal compressive failure mechanisms in notched cross-ply laminates is studied experimentally with in-situ synchrotron radiation based computed tomography. Carbon/epoxy specimens loaded monotonically in uniaxial compression exhibited a quasi-stable failure process, which was captured with computed tomography scans recorded continuously with a temporal resolutions of 2.4 seconds and a spatial resolution of 1.1 microns per voxel. A detailed chronology of the initiation and propagation of longitudinal matrix splitting cracks, in-plane and out-of-plane kink bands, shear-driven fiber failure, delamination, and transverse matrix cracks is provided with a focus on kink bands as the dominant failure mechanism. An automatic segmentation procedure is developed to identify the boundary surfaces of a kink band. The segmentation procedure enables 3-dimensional visualization of the kink band and conveys the orientation, inclination, and spatial variation of the kink band. The kink band inclination and length are examined using the segmented data revealing tunneling and spatial variations not apparent from studying the 2-dimensional section data.
    Keywords: Composite Materials
    Type: ASC Paper-49 , NF1676L-26427 , American Society for Testing and Materials (ASTM) D30 Meeting; Oct 23, 2017 - Oct 25, 2017; West Lafayette, IN; United States|American Society for Composites Technical Conference; Oct 23, 2017 - Oct 25, 2017; West Lafayette, IN; United States
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