Design and analysis of stiffened composite panels including post-buckling and collapse

Degenhardt, Richard; Kling, Alexander; Rohwer, K.; Orifici, Adrian C.; Thomson, Rodney S.

doi:10.1016/j.compstruc.2007.04.022

Cited by 70 publications

(27 citation statements)

References 18 publications

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“…It equals 90.4 kN, which is close to the end of the linear behaviour in the non linear analysis, where F cr is approximately 90 kN. Now, numerical results are concordant with the buckling mechanism [10]: from the critical buckling load, the skin local buckling causes a load redistribution in the stiffeners until collapse, and the behaviour becomes nonlinear. Thus linear models cannot be used anymore.…”

Section: Introductionsupporting

confidence: 70%

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Global behaviour of a composite stiffened panel in buckling. Part 2: Experimental investigation

Perret

Mistou

Fazzini

et al. 2012

Composite Structures

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a b s t r a c tThe present study analyses an aircraft composite fuselage structure manufactured by the Liquid Resin Infusion (LRI) process and subjected to a compressive load. LRI is based on the moulding of high performance composite parts by infusing liquid resin on dry fibres instead of prepreg fabrics or Resin Transfer Moulding (RTM). Actual industrial projects face composite integrated structure issues as a number of structures (stiffeners, . . .) are more and more integrated onto the skins of aircraft fuselage.A post-buckling test of a composite fuselage representative panel is set up, from numerical results available in previous works. Two stereo Digital Image Correlation (DIC) systems are positioned on each side of the panel, that are aimed at correlating numerical and experimental out-of-plane displacements (corresponding to the skin local buckling displacements of the panel). First, the experimental approach and the test facility are presented. A post-mortem failure analysis is then performed with the help of Non-Destructive Techniques (NDT). X-ray Computed Tomography (CT) measurements and ultrasonic testing (US) techniques are able to explain the failure mechanisms that occured during this post-buckling test. Numerical results are validated by the experimental results.

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Section: Introductionsupporting

confidence: 70%

“…10a. The experimental critical buckling load corresponds to skin local buckling at approximately 90 kN, which is concordant with the buckling mechanism [10]. Five symmetrical waves are formed on each panel skin as it has also been predicted by the linear buckling analysis (see Fig.…”

Section: Skin Critical Buckling Loadsupporting

confidence: 50%

Global behaviour of a composite stiffened panel in buckling. Part 2: Experimental investigation

Perret

Mistou

Fazzini

et al. 2012

Composite Structures

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“…It is followed by local skin buckling which is mixed with global buckling. This sequence is not concordant with the given description of the buckling mechanism [21]. That can be explained first by the flat shape of the studied panel which leads the panel to buckle globally on the stiffener side, plus no gliding conditions on the longitudinal edges of the panels.…”

Section: Geometric Non Linear Resultsmentioning

confidence: 69%

“…Buckling load evolution on load shortening curves can be ideally described with three marked load levels [21], the first local buckling load being the skin buckling between the stiffeners, then the first global buckling load being governed by the stiffener load carriage capacity, finally leading to the collapse load which is the highest load. This sequence of mechanisms has been originally elaborated for describing the global behaviour of the debonding of prepregs composite panels made of co-bonding structures.…”

Section: Geometric Non Linear Resultsmentioning

confidence: 99%

Global behaviour of a composite stiffened panel in buckling. Part 1: Numerical modelling

Perret¹,

Mistou²,

Fazzini³

2011

Composite Structures

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a b s t r a c tThe present study analyses an aircraft composite fuselage structure manufactured by the Liquid Resin Infusion (LRI) process and subjected to a compressive load. LRI is based on the moulding of high performance composite parts by infusing liquid resin on dry fibres instead of prepreg fabrics or Resin Transfer Moulding (RTM). Actual industrial projects face composite integrated structure issues as a number of structures (stiffeners, . . .) are more and more integrated onto the skins of aircraft fuselage. A representative panel of a composite fuselage to be tested in buckling is studied numerically. This paper studies which of the real behaviours of the integrated structures are to be observed during this test. Numerical models are studied at a global scale of the composite stiffened panel. Linear and non linear analyses are conducted. The Tsai-Wu criterion with a progressive failure analysis is implemented, to describe the global behaviour of the panel up to collapse. Also, three stiffener connection methods are compared at the intersection between two types of integrated structures. Load shortening curves permit to estimate the expected load and displacements.

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“…For the Design 1 there is a large postbuckling region, even after the first global stringer buckling and the stringer buckling starts in the middle of the panel. DLR's experience in the designing of panels within the projects POSICOSS and COCOMAT is explained in more detail in [6].…”

Section: B Workpackagementioning

confidence: 99%

Improved Design Scenario for Composite Airframe Structures

Degenhardt¹,

Teßmer²

2007

48th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference

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European aircraft industry demands for reduced development and operating costs, by 20% and 50% in the short and long term, respectively. Contributions to this aim are provided by the completed project POSICOSS and the running follow-up project COCOMAT, both supported by the European Commission. As an important contribution to cost reduction a decrease in structural weight can be reached by exploiting considerable reserves in primary fibre composite fuselage structures through an accurate and reliable simulation of postbuckling and collapse. The POSICOSS team developed fast procedures for postbuckling analysis of fibre composite stiffened panels, created comprehensive experimental data bases and derived design guidelines. COCOMAT builds up on the POSICOSS results and considers in addition the simulation of collapse by taking degradation into account. The main objective is a future design scenario for composite stiffened panels which allows the exploiting of considerable reserves in primary fibre composite fuselage structures. The results comprise an extended experimental data base, degradation models, improved certification and design tools as well as design guidelines. The paper deals with the main objectives of the project COCOMAT, the general status of the progress as well as DLR's first results.

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Design and analysis of stiffened composite panels including post-buckling and collapse

Cited by 70 publications

References 18 publications

Global behaviour of a composite stiffened panel in buckling. Part 2: Experimental investigation

Global behaviour of a composite stiffened panel in buckling. Part 2: Experimental investigation

Global behaviour of a composite stiffened panel in buckling. Part 1: Numerical modelling

Improved Design Scenario for Composite Airframe Structures

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