A Finite Element Methodology for Analysing Degradation and Collapse in Postbuckling Composite Aerospace Structures

Orifici, Adrian C.; Thomson, Rodney S.; Degenhardt, Richard; Bisagni, Chiara; Bayandor, Javid

doi:10.1177/0021998309345294

Cited by 51 publications

(31 citation statements)

References 31 publications

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“…Figure 5 and Table 1 present the detailed geometry of the T-joint where "t_ply", "t_st", "t_s" are the thicknesses of the single ply, stringer and skin, respectively, "L_st" and "L_s" are the length of the stringer and the skin, respectively, "h_st" is the height of the stringer. IM7/8552 composite material elastic and interface properties for skin and stringer legs and FM300 adhesive material elastic and interface properties for filler and bonding lines are used which are presented in Table 2 and Table 3 [2,24]. 0° ply direction is defined to be in the direction of the stringer, namely out of plane direction in this case.…”

Section: T-joint Geometry and The Finite Element Modelmentioning

confidence: 99%

“…Under various loading conditions (mechanical, buckling, cyclic, hydrothermal, low velocity impact, bird strike, underwater explosion, lightning, etc. [2,3,4,5,14,21]), interlaminar normal and shear stresses are generated between the composite plies in these critical regions which cause delaminations/debonds. Delamination/debond reduces the stiffness and strength of the structure as the delaminated area gets larger.…”

Section: Introductionmentioning

confidence: 99%

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Delamination-Debond Behaviour of Composite T- Joints in Wind Turbine Blades

Gülaşık¹,

Çöker²

2014

J. Phys.: Conf. Ser.

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Abstract. Wind turbine industry utilizes composite materials in turbine blade structural designs because of their high strength/stiffness to weight ratio. T-joint is one of the design configurations of composite wind turbine blades. T-joints consist of a skin panel and a stiffener co-bonded or co-cured together with a filler material between them. T-joints are prone to delaminations between skin/stiffener plies and debonds between skin-stiffener-filler interfaces. In this study, delamination/debond behavior of a co-bonded composite T-joint is investigated under 0° pull load condition by 2D finite element method. Using Abaqus ® commercial FE software, zero-thickness cohesive elements are used to simulate delamination/debond in ply interfaces and bonding lines. Pulling load at 0° is applied and load-displacement behavior and failure scenario are observed. The failure sequence consists of debonding of filler/stringer interface during one load drop followed by a second drop in which the 2 nd filler/stringer debonds, filler/skin debonding and skin delamination leading to total loss of load carrying capacity. This type of failure initiation has been observed widely in the literature. When the debond strength is increased 30%, failure pattern is found to change in addition to increasing the load capacity by 200% before total loss of loading carrying capacity occurs. Failure initiation and propagation behavior, initial and max failure loads and stress fields are affected by the property change. In all cases mixed-mode crack tip loading is observed in the failure initiation and propagation stages. In this paper, the detailed delamination/debonding history in T-joints is predicted with cohesive elements for the first time.

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Section: T-joint Geometry and The Finite Element Modelmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

Delamination-Debond Behaviour of Composite T- Joints in Wind Turbine Blades

Gülaşık¹,

Çöker²

2014

J. Phys.: Conf. Ser.

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“…• Two pre-damaged multi-stiffener D1 panels manufactured at Aernnova and tested at DLR, the first in which skin-stiffener debonds were introduced using cyclic loading into the postbuckling region, 10 and another that used Teflon-generated skin-stiffener debonds. 7 In general, the comparison with all experimental results demonstrated that the analysis tool was able to give accurate predictions of the critical damage mechanisms, panel behaviour and collapse load.…”

Section: Comparison With Experimentsmentioning

confidence: 99%

An Analysis Tool for Design and Certification of Postbuckling Composite Aerospace Structures

Orifici

Thomson²,

Degenhardt

et al. 2010

Int. J. Str. Stab. Dyn.

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In aerospace, carbon fibre-reinforced polymer (CFRP) materials and postbuckling skin-stiffened structures are key technologies that have been used to improve structural efficiency. However, the application of composite postbuckling structures in aircraft has been limited as today's analysis tools cannot accurately predict structural collapse in compression. In this work, a finite element analysis tool for design and certification of aerospace structures is presented, which predicts collapse by taking the critical damage mechanisms into account. The tool incorporates a global-local analysis technique for predicting interlaminar damage initiation, and degradation models to capture the growth of a pre-existing interlaminar damage region, such as a delamination or skin-stiffener debond, and in-plane ply damage mechanisms such as fibre fracture and matrix cracking. The analysis tool has been applied to single-and multi-stiffener fuselage-representative composite panels, in both intact and pre-damaged configurations. This has been performed in a design context, in which panel configurations are selected based on their suitability for experimental testing, and in an analysis context for comparison with experimental results as representative of aircraft certification studies. For all cases, the tool was capable of accurately capturing the key damage mechanisms contributing to final structural collapse, and suitable for the design of next-generation composite aerospace structures. Analysis ToolThe finite element (FE) analysis tool developed was focused on predicting the collapse of stiffened composite structures in compression by capturing the effects of the critical damage mechanisms. The approach contains several aspects: predicting initiation of interlaminar damage in intact structures; capturing in-plane degradation; and, capturing the propagation of a pre-existing interlaminar damage region. The analysis tool was implemented into MSC.Marc v2005r3 (Marc) by a combination of user subroutines, and supported by utilities for pre-and post-processing within a user-friendly menu system developed in MSC.Patran. 1 This work was conducted as part of the European Commission Project COCOMAT, which focused on exploiting the large strength reserves of composite aerospace structures through a more accurate prediction of collapse. 2The approach for predicting the initiation of interlaminar damage in the skin-stiffener interface was based on a globallocal technique. In this technique, a global shell model of the full structure was used to determine the deformation field of the entire structure, which was then input as boundary conditions on a local three-dimensional (3D) brick model of a skin-stiffener interface. In the local model the strength-based "degenerated Tsai" criterion was monitored at all elements in order to predict the initiation of delamination or skin-stiffener separation. Failure was deemed to occur when the average of all integration point values in an element satisfied this criterion. By modifying the locat...

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“…A finite element analysis tool for design and certification of aerospace structures is presented by Orifici et al [7], which incorporates a global-local analysis technique for predicting interlaminar damage initiation, and degradation models to capture the growth of a pre-existing interlaminar damage region, such as a delamination or skin-stiffener debond, and in-plane ply damage mechanisms such as fibre fracture and matrix cracking.…”

Section: Introductionmentioning

confidence: 99%

The post-buckling response of stiffened CFRP panels using a single-stringer compression specimen (SSCS): a numerical investigation

Sadiq¹,

Qing²

2015

Computational Methods and Experimental Measurements XVII

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The safe exploitation of the post-buckling region of CFRP panels in aerospace applications is of prime interest here. Due to the long computational time of finite element analysis for the post-buckling response of CFRP panels, there is a need to devise some efficient procedures. Skin/stringer debonding is primarily responsible for the collapse of the stiffened panels under compressive loading. In this study a simplified approach based on a single-stringer compression specimen (SSCS) has been employed to gain an insight into the post-buckling behaviour of CFRP panels under axial compression. A progressive damage model is used to get more accurate numerical results of the post-buckling response of SSCS. The outcomes of the study are compared with the published results and a fair agreement is found. In addition, a comparative study of open-and closed-section stringer stiffened panels based on a global-local approach for predicting pre-buckling, post-buckling and mode shapes is carried out. As the simplified model is used in this study so detailed damage models can be constructed to account for all damage modes: matrix cracking, fibre kinking, fibre fracture, and delamination. The study can be employed as a global-local approach to get the post-buckling response of CFRP stiffened panels. Furthermore, it also provides an efficient mean for the design optimization of CFRP structures in the post-buckling regime.

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A Finite Element Methodology for Analysing Degradation and Collapse in Postbuckling Composite Aerospace Structures

Cited by 51 publications

References 31 publications

Delamination-Debond Behaviour of Composite T- Joints in Wind Turbine Blades

Delamination-Debond Behaviour of Composite T- Joints in Wind Turbine Blades

An Analysis Tool for Design and Certification of Postbuckling Composite Aerospace Structures

The post-buckling response of stiffened CFRP panels using a single-stringer compression specimen (SSCS): a numerical investigation

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