The Experts below are selected from a list of 213 Experts worldwide ranked by ideXlab platform

Dawn C. Jegley - One of the best experts on this subject based on the ideXlab platform.

  • Testing a Multi-bay Box Subjected to Combined Loads
    Conference Proceedings of the Society for Experimental Mechanics Series, 2016
    Co-Authors: Marshall Rouse, Dawn C. Jegley
    Abstract:

    The COmbined Loads Test System (COLTS) facility at NASA Langley Research Center provides a test capability to help develop validated structures technologies. The test machine was Design to accommodate a range of fuselage structures and wing sections and subject them to both quasistatic and cyclic Loading conditions. The COLTS facility is capable of testing fuselage barrels up to 4.6 m in diameter and 13.7 m long with combined mechanical, internal pressure, and thermal Loads. The COLTS facility is currently being prepared to conduct a combined mechanical and pressure Loading for a multi-bay pressure box to experimentally verify the structural performance of a composite structure which is 9.1 m long and representative of a section of a hybrid wing body fuselage section in support of the Environmentally Responsible Aviation Project at NASA. This paper describes development of the multi-bay pressure box test using the COLTS facility. The multi-bay test article will be subjected to mechanical Loads and internal pressure Loads up to Design Ultimate Load. Mechanical and pressure Loads will be applied independently in some tests and simultaneously in others.

  • Preparation for Testing a Multi-Bay Box Subjected to Combined Loads
    2015
    Co-Authors: Marshall Rouse, Dawn C. Jegley
    Abstract:

    The COmbined Loads Test System (COLTS) facility at NASA Langley Research Center provides a test capability to help develop validated structures technologies. The test machine was Design to accommodate a range of fuselage structures and wing sections and subject them to both quasistatic and cyclic Loading conditions. The COLTS facility is capable of testing fuselage barrels up to 4.6 m in diameter and 13.7 m long with combined mechanical, internal pressure, and thermal Loads. The COLTS facility is currently being prepared to conduct a combined mechanical and pressure Loading for a multi-bay pressure box to experimentally verify the structural performance of a composite structure which is 9.1 meters long and representative of a section of a hybrid wing body fuselage section in support of the Environmentally Responsible Aviation Project at NASA. This paper describes development of the multi-bay pressure box test using the COLTS facility. The multi-bay test article will be subjected to mechanical Loads and internal pressure Loads up to Design Ultimate Load. Mechanical and pressure Loads will be applied independently in some tests and simultaneously in others.

  • Structural Efficiency and Behavior of Pristine and Notched Stitched Structure
    2011
    Co-Authors: Dawn C. Jegley
    Abstract:

    Two driving factors in aircraft panel Design are structural efficiency and response to in-service damage. Stitching through the thickness can improve both of these considerations. Combining stitching with a post-buckling Design approach can provide additional improvements. The buckling behavior of stitched structure is considered since lighter structures can be achieved if local skin buckling is allowed to occur at less than Design Ultimate Load. Through-the-thickness stitching can suppress delamination between skin and flange, thereby allowing the structure to reliably carry Load into the postbuckling range. Hat-stiffened and rod-stiffened panels in which the skin and flanges were stitched together through-the-thickness prior to curing are considered through experiment and analysis. In both types of panels no mechanical fasteners were used for the assembly. Specimens were Loaded to failure in axial compression. In this study all specimens buckled in the skin between the stiffeners and continued to carry Load. In addition, the behavior of panels with a severed stringer or notch are considered. Failure Loads and strain distributions in the notched panel are compared to those in the unnotched panel.

  • Structural Efficiency of Stitched Rod-Stiffened Composite Panels with Stiffener Crippling
    49th AIAA ASME ASCE AHS ASC Structures Structural Dynamics and Materials Conference <br> 16th AIAA ASME AHS Adaptive Structures Conference<br, 2008
    Co-Authors: Dawn C. Jegley, Alexander Velicki, Daniel A. Hansen
    Abstract:

    The structural efficiency of rod-stiffened stitched specimens is evaluated to determine their weight saving potential if the stiffeners were allowed to buckle at less than or equal to Design Ultimate Load. Analytical and experimental results from rod-stiffened and blade-stiffened single-stiffener specimens are presented. In both cases, skin and flanges were stitched together through-the-thickness prior to curing. No mechanical fasteners were used for the assembly. Specimens were Loaded to failure in axial compression. Failure modes are discussed. Finite element and experimental results agree for the response of the structures. For some specimen configurations, improved structural efficiency can be obtained by allowing stiffeners to buckle at Design limit Load rather than requiring that buckling not occur prior to Design Ultimate Load. In addition, through-the-thickness stitching can change the failure mechanism by suppressing delamination between skin and flange. A parametric study is presented herein which describes the possible weight savings with this approach.

  • Structural Efficiency of Stitched Composite Panels with Stiffener Crippling
    Journal of Aircraft, 2005
    Co-Authors: Dawn C. Jegley
    Abstract:

    The structural efficiency of blade-stiffened stitched specimens is compared to determine their weight-saving potential if blades were allowed to buckle at less than or equal to Design Ultimate Load. Analytical and experimental results from four configurations of crippling specimens are presented. Specimen skin and blades were held together with through-the-thickness stitches prior to curing. No mechanical fasteners were used for the assembly. Tests were conducted with and without low-speed impact damage. Failure modes are discussed. Finite element and experimental results agree for the response of the structures. For some specimen configurations, improved structural efficiency can be obtained by allowing stiffeners to buckle at Design limit Load rather than requiring that buckling not occur prior to Design Ultimate Load. A parametric study is presented herein, which describes the possible weight savings with this approach.

Chun H. Wang - One of the best experts on this subject based on the ideXlab platform.

  • Design and optimization of scarf repairs
    Bonded Joints and Repairs to Composite Airframe Structures, 2016
    Co-Authors: Chun H. Wang, Chun H. Wang, Cong N. Duong
    Abstract:

    Several techniques to reduce the size of flush repairs are presented in this chapter. The strength requirement akin to the safe-life approach is first discussed by focusing on the residual strength of a composite structure containing a scarfed hole. Selection of the scarf angle, which directly affects the size of repair, needs to meet both the Design limit Load and Design Ultimate Load requirements. Reduction in repair size can be further accomplished by optimizing the shape of repairs and the incorporation of a structural doubler. Design methods are presented together with computational and experimental verifications.

  • Effects of bondline flaws on the damage tolerance of composite scarf joints
    Composites Part A-applied Science and Manufacturing, 2013
    Co-Authors: S. Georgiadis, Adrian C. Orifici, Chun H. Wang
    Abstract:

    Scarf repairs to aircraft structures need to sustain Design Ultimate Load in the presence of flaws due to manufacturing and impact by foreign objects, in order to demonstrate compliance with airworthiness regulations. This paper presents an investigation into the effect of disbonds on the Load-carrying capacity of adhesively bonded scarf joints. Experiments were conducted on scarf joints containing disbonds of varying lengths. The results showed that the Load-carrying capacity of scarf joints decreases with the size of the bondline flaw at a faster rate than the reduction in the effective bond area. Fractographic analysis showed that the fracture occurred in the composite matrix adjacent to the adhesive-composite interface, at a distance equal to a small fraction of ply thickness. Computational analyses using the virtual crack closure technique (VCCT) and the cohesive zone model (CZM) confirmed these experimental observations: model predictions using composite material properties were in better correlation with experimental results than those using adhesive properties. Furthermore, CZM is capable of predicting the effects of flaws of all sized being considered, while the VCCT model is only applicable to joints containing flaws greater than a certain size.

  • Effects of bondline flaws on the damage tolerance of composite scarf joints
    Composites Part A: Applied Science and Manufacturing, 2013
    Co-Authors: J.y. Goh, Adrian C. Orifici, S. Georgiadis, Chun H. Wang
    Abstract:

    Scarf repairs to aircraft structures need to sustain Design Ultimate Load in the presence of flaws due to manufacturing and impact by foreign objects, in order to demonstrate compliance with airworthiness regulations. This paper presents an investigation into the effect of disbonds on the Load-carrying capacity of adhesively bonded scarf joints. Experiments were conducted on scarf joints containing disbonds of varying lengths. The results showed that the Load-carrying capacity of scarf joints decreases with the size of the bondline flaw at a faster rate than the reduction in the effective bond area. Fractographic analysis showed that the fracture occurred in the composite matrix adjacent to the adhesive-composite interface, at a distance equal to a small fraction of ply thickness. Computational analyses using the virtual crack closure technique (VCCT) and the cohesive zone model (CZM) confirmed these experimental observations: model predictions using composite material properties were in better correlation with experimental results than those using adhesive properties. Furthermore, CZM is capable of predicting the effects of flaws of all sized being considered, while the VCCT model is only applicable to joints containing flaws greater than a certain size.

S. Georgiadis - One of the best experts on this subject based on the ideXlab platform.

  • Effects of bondline flaws on the damage tolerance of composite scarf joints
    Composites Part A-applied Science and Manufacturing, 2013
    Co-Authors: S. Georgiadis, Adrian C. Orifici, Chun H. Wang
    Abstract:

    Scarf repairs to aircraft structures need to sustain Design Ultimate Load in the presence of flaws due to manufacturing and impact by foreign objects, in order to demonstrate compliance with airworthiness regulations. This paper presents an investigation into the effect of disbonds on the Load-carrying capacity of adhesively bonded scarf joints. Experiments were conducted on scarf joints containing disbonds of varying lengths. The results showed that the Load-carrying capacity of scarf joints decreases with the size of the bondline flaw at a faster rate than the reduction in the effective bond area. Fractographic analysis showed that the fracture occurred in the composite matrix adjacent to the adhesive-composite interface, at a distance equal to a small fraction of ply thickness. Computational analyses using the virtual crack closure technique (VCCT) and the cohesive zone model (CZM) confirmed these experimental observations: model predictions using composite material properties were in better correlation with experimental results than those using adhesive properties. Furthermore, CZM is capable of predicting the effects of flaws of all sized being considered, while the VCCT model is only applicable to joints containing flaws greater than a certain size.

  • Effects of bondline flaws on the damage tolerance of composite scarf joints
    Composites Part A: Applied Science and Manufacturing, 2013
    Co-Authors: J.y. Goh, Adrian C. Orifici, S. Georgiadis, Chun H. Wang
    Abstract:

    Scarf repairs to aircraft structures need to sustain Design Ultimate Load in the presence of flaws due to manufacturing and impact by foreign objects, in order to demonstrate compliance with airworthiness regulations. This paper presents an investigation into the effect of disbonds on the Load-carrying capacity of adhesively bonded scarf joints. Experiments were conducted on scarf joints containing disbonds of varying lengths. The results showed that the Load-carrying capacity of scarf joints decreases with the size of the bondline flaw at a faster rate than the reduction in the effective bond area. Fractographic analysis showed that the fracture occurred in the composite matrix adjacent to the adhesive-composite interface, at a distance equal to a small fraction of ply thickness. Computational analyses using the virtual crack closure technique (VCCT) and the cohesive zone model (CZM) confirmed these experimental observations: model predictions using composite material properties were in better correlation with experimental results than those using adhesive properties. Furthermore, CZM is capable of predicting the effects of flaws of all sized being considered, while the VCCT model is only applicable to joints containing flaws greater than a certain size.

Andrew E. Lovejoy - One of the best experts on this subject based on the ideXlab platform.

  • Structural Response and Failure of a Full-Scale Stitched Graphite-Epoxy Wing
    Journal of Aircraft, 2003
    Co-Authors: Dawn C. Jegley, Harold G. Bush, Andrew E. Lovejoy
    Abstract:

    Analytical and experimental results of the test for an all-composite full-scale wing box are presented. The wing box is representative of a section of a 220-passenger commercial transport aircraft wing box and was Designed and constructed by The Boeing Company as part of the NASA Advanced Subsonics Technology (AST) program. The semi-span wing was fabricated from a graphite-epoxy material system with cover panels and spars held together using Kevlar stitches through the thickness. No mechanical fasteners were used to hold the stiffeners to the skin of the cover panels. Tests were conducted with and without low-speed impact damage, discrete source damage and repairs. Up-bending down-bending and brake roll Loading conditions were applied. The structure with nonvisible impact damage carried 97% of Design Ultimate Load prior to failure through a lower cover panel access hole. Finite element and experimental results agree for the global response of the structure.

  • Evaluation of the Structural Response and Failure of a Full-Scale Stitched Graphite-Epoxy Wing
    2001
    Co-Authors: Dawn C. Jegley, Harold G. Bush, Andrew E. Lovejoy
    Abstract:

    Analytical and experimental results for an all-composite full-scale wing box are presented. The wing box is representative of a section of a 220-passenger commercial transport aircraft wing box and was Designed and constructed by The Boeing Company as part of the NASA Advanced Subsonics Technology (AST) program. The semi-span wing was fabricated from a graphite-epoxy material system with cover panels and spars held together using Kevlar stitches through the thickness. No mechanical fasteners were used to hold the stiffeners to the skin of the cover panels. Tests were conducted with and without low-speed impact damage, discrete source damage and repairs. Upbending, down-bending and brake roll Loading conditions were applied. The structure with nonvisible impact damage carried 97% of Design Ultimate Load prior to failure through a lower cover panel access hole. Finite element and experimental results agree for the global response of the structure.

Adrian C. Orifici - One of the best experts on this subject based on the ideXlab platform.

  • Effects of bondline flaws on the damage tolerance of composite scarf joints
    Composites Part A-applied Science and Manufacturing, 2013
    Co-Authors: S. Georgiadis, Adrian C. Orifici, Chun H. Wang
    Abstract:

    Scarf repairs to aircraft structures need to sustain Design Ultimate Load in the presence of flaws due to manufacturing and impact by foreign objects, in order to demonstrate compliance with airworthiness regulations. This paper presents an investigation into the effect of disbonds on the Load-carrying capacity of adhesively bonded scarf joints. Experiments were conducted on scarf joints containing disbonds of varying lengths. The results showed that the Load-carrying capacity of scarf joints decreases with the size of the bondline flaw at a faster rate than the reduction in the effective bond area. Fractographic analysis showed that the fracture occurred in the composite matrix adjacent to the adhesive-composite interface, at a distance equal to a small fraction of ply thickness. Computational analyses using the virtual crack closure technique (VCCT) and the cohesive zone model (CZM) confirmed these experimental observations: model predictions using composite material properties were in better correlation with experimental results than those using adhesive properties. Furthermore, CZM is capable of predicting the effects of flaws of all sized being considered, while the VCCT model is only applicable to joints containing flaws greater than a certain size.

  • Effects of bondline flaws on the damage tolerance of composite scarf joints
    Composites Part A: Applied Science and Manufacturing, 2013
    Co-Authors: J.y. Goh, Adrian C. Orifici, S. Georgiadis, Chun H. Wang
    Abstract:

    Scarf repairs to aircraft structures need to sustain Design Ultimate Load in the presence of flaws due to manufacturing and impact by foreign objects, in order to demonstrate compliance with airworthiness regulations. This paper presents an investigation into the effect of disbonds on the Load-carrying capacity of adhesively bonded scarf joints. Experiments were conducted on scarf joints containing disbonds of varying lengths. The results showed that the Load-carrying capacity of scarf joints decreases with the size of the bondline flaw at a faster rate than the reduction in the effective bond area. Fractographic analysis showed that the fracture occurred in the composite matrix adjacent to the adhesive-composite interface, at a distance equal to a small fraction of ply thickness. Computational analyses using the virtual crack closure technique (VCCT) and the cohesive zone model (CZM) confirmed these experimental observations: model predictions using composite material properties were in better correlation with experimental results than those using adhesive properties. Furthermore, CZM is capable of predicting the effects of flaws of all sized being considered, while the VCCT model is only applicable to joints containing flaws greater than a certain size.