| Auteur | Rechercher : Laliberte, J. F.1; Rechercher : Poon, C.1; Rechercher : Fahr, A.1; Rechercher : Straznicky, P. V.1 |
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| Affiliation | - Conseil national de recherches Canada. Institut de recherche aérospatiale du CNRC
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| Format | Texte, Résumé |
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| Conférence | Second International Conference on Fatigue of Composites, June 4-7, 2000, Williamsburg, VA, USA |
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| Résumé | Fibre-metal laminates (FMLs) represent a significant evolution in aircraft structural materials. These materials are hybrid laminates ofmetal sheets with fibre-reinforced polymer layers. This provides a combination oflow density, fatigue damage growth resistance., impact damage tolerance and high static strength that is extremely well suited for aerospace structures. Currently FMLs are experiencing increasing use as secondary airframe materials in applications such as control surfaces, cargo hold floors and doors. They are also being considered for numerous primary structural (fuselage, empennage, wings) applications in aircraft under development or in production.
Full characterization of the properties of these laminates is required before they can be certified for primary airframe applications. Therefore, the Fibre-Metal Laminate Durability Project was established as a jointly funded project between Bombardier Aerospace, the Institute for Aerospace Research (IAR) of the National Research Council Canada (NRC) and Carleton University (Ottawa). The aim ofthis project is to characterize the impact resistance, post-impact residual strength and multi-site damage growth behaviour of GLARE (GLAss REinforced) fibre-metal laminates.
Following the impact event there is a plastically deformed dent in FML panels that acts as a stress concentration and as an initiation site for damage growth. Due to the residual deformation in the panels the growth mechanism is significantly more complicated than that reported by other researchers in flat FML panels with holes or notches. One particular laminate of interest is GLARE-4-2/1 which consists of two outer layers of 2024-T3 aluminum and three unidirectional plies of S2-glass reinforced epcxy in the following lay-up [2024- T3/(0°/90°/0°)/2024-T3]. This lay-up is typical or what would be used in pressurized fuselage applications where a biaxial stress state exists.
Non-destructive inspection techniques such as a penetrant enhanced x-ray process developed at IAR and ultrasonic C-scan were employed to characterize the fatigue damage growth at various stages. Initially, crack-growth actually proceeds parallel to the loading axis from the centre of the dent. This only proceeds until the crack reaches the edge of the dented region. Cracks also develop perpendicular to the loading direction and are associated with large regions of delamination between the aluminum layers and the fibre layers. The fibre layers bridge the crack tip and contribute to a residual fatigue life significantly higher than that observed in monolithic aluminum specimens. To better understand the crack growth mechanisms the stress state in the dented region needs to be examined in detail. The results of the experimental testing and a proposed modeling approach will be reported in this paper. |
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| Date de publication | 2000-06 |
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| Maison d’édition | The Conference |
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| Dans | |
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| Langue | anglais |
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| Publications évaluées par des pairs | Oui |
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| Numéro du CNRC | SMPL-1999-0090 |
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| Numéro NPARC | 8930410 |
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| Exporter la notice | Exporter en format RIS |
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| Signaler une correction | Signaler une correction (s'ouvre dans un nouvel onglet) |
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| Identificateur de l’enregistrement | c8636c6c-543a-436c-b046-e6bb9b705d9a |
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| Enregistrement créé | 2009-04-23 |
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| Enregistrement modifié | 2023-02-23 |
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