<abstract xmlns="http://www.w3.org/1999/xhtml"><p>Damage tolerance of composite aircraft structure is one of the main areas of research, important when a new product is being developed. There are a number of variables, such as damage characteristics (dent depth, delamination area) and loading parameters (load type, amplitude of cyclic loading, load sequence) that need to be investigated experimentally [1]. These tests of composite materials are usually performed at an element level and are carried out in order to validate the analytical model, developed to predict the full-scale component’s behaviour. The paper presents the results of compression testing of the [36/55/9] carbon fibre/epoxy laminate, manufactured with the Automated Fibre Placement technology (AFP) and subjected to static and fatigue loads. The laminate compression loading mode was achieved through sandwich 4-point flexure. At the stage of fatigue testing, two parameters were investigated: the damage size, simulated by the hole diameter and the fatigue load level. Based on the test results, the laminate fatigue load limit equal to 75% of the OHC failure load was evaluated. By collating the static and fatigue tests results, the damage tolerance characteristic of the considered laminate was created.</p></abstract>

Damage tolerance of composite aircraft structure is one of the main areas of research, important when a new product is being developed. There are a number of variables, such as damage characteristics (dent depth, delamination area) and loading parameters (load type, amplitude of cyclic loading, load sequence) that need to be investigated experimentally [1]. These tests of composite materials are usually performed at an element level and are carried out in order to validate the analytical model, developed to predict the full-scale component’s behaviour. The paper presents the results of compression testing of the [36/55/9] carbon fibre/epoxy laminate, manufactured with the Automated Fibre Placement technology (AFP) and subjected to static and fatigue loads. The laminate compression loading mode was achieved through sandwich 4-point flexure. At the stage of fatigue testing, two parameters were investigated: the damage size, simulated by the hole diameter and the fatigue load level. Based on the test results, the laminate fatigue load limit equal to 75% of the OHC failure load was evaluated. By collating the static and fatigue tests results, the damage tolerance characteristic of the considered laminate was created.

Influence of the Fatigue Load Level and the Hole Diameter on the Laminate Structure’s Fatigue Performance

Fatigue of Aircraft Structures

This work is licensed under the Creative Commons Attribution-NonCommercial-NoDerivatives 3.0 License.

Damage tolerance of composite aircraft structure is one of the main areas of research, important when a new product is being developed. There are a number of...

Influence of the Fatigue Load Level and the Hole Diameter on the Laminate Structure’s Fatigue Performance

Online veröffentlicht: 28. Juni 2019

Seitenbereich: 21 - 30

DOI: https://doi.org/10.2478/fas-2018-0002

Schlüsselwörter
composite, fatigue, damage

© 2019 Małgorzata Zalewska, published by Sciendo

This work is licensed under the Creative Commons Attribution-NonCommercial-NoDerivatives 3.0 License.

Influence of the Fatigue Load Level and the Hole Diameter on the Laminate Structure’s Fatigue Performance

Online veröffentlicht: 28. Juni 2019

Seitenbereich: 21 - 30

DOI: https://doi.org/10.2478/fas-2018-0002

Schlüsselwörtercomposite, fatigue, damage

© 2019 Małgorzata Zalewska, published by Sciendo

This work is licensed under the Creative Commons Attribution-NonCommercial-NoDerivatives 3.0 License.

Schlüsselwörter
composite, fatigue, damage