The use of damage as a design parameter for postbuckling composite aerospace structures

Advanced fibre-reinforced polymer composites have seen a rapid increase in use in aircraft structures in recent years due their high specific strength and stiffness, amongst other properties. The use of postbuckling design, where lightweight structures are designed to operate safely at loads in excess of buckling loads, has been applied to metals for decades to design highly efficient structures. However, to date, the application of postbuckling design in composite structures has been limited, as today’s analysis tools are not capable of representing the damage mechanisms that lead to structural collapse of composites in compression. The currently running four-year European Commission Project COCOMAT [1] is addressing this issue, and aims to exploit the large strength reserves of composite aerospace structures through a more accurate prediction of collapse. A methodology has been developed to analyse the collapse of composite structures that is focused on capturing the critical damage mechanisms. One aspect of the methodology is a global-local analysis technique that uses a strength criterion to predict the initiation of interlaminar damage in intact structures. Another aspect of the approach was developed for representing the growth of a pre-existing interlaminar damage region, and is based on applying multi-point constraints in the skin-stiffener interface that are controlled using fracture mechanics calculations. A separate degradation model was also included to model the in-plane ply damage mechanisms of fibre fracture, matrix cracking and fibre-matrix shear that uses a progressive failure approach. The complete analysis methodology was implemented in MSC.Marc v2005r3 using several user subroutines, and has been validated with a range of experimental tests, including fracture mechanics coupons [2], single-stiffener specimens [3] and multi-stiffener curved panels [4]. The developed methodology was used to design and analyse fuselage-representative composite panels in various pre-damaged configurations. Two panel designs were investigated, D1 and D2, which both consisted of a curved skin adhesively bonded to blade-shaped stiffeners. For the D1 panel, the pre-damage applied was a full-width skin-stiffener debond created using a Teflon insert in the adhesive layer, whilst the D2 panel was investigated with Barely Visible Impact Damage (BVID). For both panels, parametric studies were conducted using the developed methodology in order to recommend a damaged configuration suitable for experimental testing. For the D1 panel, a 100 mm length debond was selected, and the location of the damage was investigated, whilst for the D2 panel both the location and the representation of damage was varied. Based on these parametric studies, two pre-damaged configurations of the D1 panel and one pre-damaged D2 configuration were selected for experimental testing. The selected pre-damaged configurations were manufactured by Aernnova Engineering Solutions and manufactured at the Institute of Composite Structures and Adaptive Systems at the German Aerospace Center (DLR) as part of the COCOMAT project. Following manufacture, panel quality was inspected with ultrasonic and thermographic scanning and panel imperfection data was measured using the three-dimensional (3D) optical measurement system ATOS. During the test, measurements were taken using displacement transducers, strain gauges, the 3D optical measuring system ARAMIS, and optical lock-in thermography. Under compression, the panels developed a range of buckling mode shapes, and the progression of damage was monitored leading to structural collapse. In comparison with the experimental results, the analysis methodology was shown to give accurate predictions of the load-carrying behaviour, damage development and collapse load of both panels. The results demonstrated the capability of the developed tool to capture the critical damage mechanisms leading to collapse in composite structures. The advanced analysis methodology also allowed for damage to be used as a design parameter in postbuckling structures, either in the comparative analysis context of a design procedure, to assess the damage tolerance of a design, or as pre- and post-test simulations of intact and pre-damaged structures. More broadly, the results demonstrated the potential of postbuckling composite structures, and the large strength reserve available in the postbuckling region. The success of the developed analysis methodology and the potential of postbuckling composite structures have application for the next generation of lightweight aerospace structures

A finite element methodology for analysing degradation and collapse in postbuckling composite aerospace structures

A methodology for analysing the degradation and collapse in postbuckling composite structures is proposed. One aspect of the methodology predicts the initiation of interlaminar damage using a strength criterion applied with a global-local analysis technique. A separate approach represents the growth of a pre-existing interlaminar damage region with user-defined multi-point constraints that are controlled based on the Virtual Crack Closure Technique. Another aspect of the approach is a degradation model for in-plane ply damage mechanisms of fiber fracture, matrix cracking, and fiber-matrix shear. The complete analysis methodology was compared to experimental results for two fuselage-representative composite panels tested to collapse. For both panels, the behavior and structural collapse were accurately captured, and the analysis methodology provided detailed information on the development and interaction of the various damage mechanisms

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

The 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

Control of composite material crack branching for improved fracture toughness

An experimental investigation was conducted to develop techniques for controlling crack branching in composite laminates. Double cantilever beam specimens were tested to study mode I dominated crack growth. Embedded flaws were generated using ply gaps and strips of non-stick film, both individually and in combination as a "branch flaw". Crack branching in 0° plies was generated in an inconsistent manner using ply gaps, but in a consistent manner using branch flaws. Branching through 90° plies occurred automatically due to their orientation, and could be further controlled using embedded delamination flaws. Crack branching in 45° plies was more complex, but could be controlled using ply gaps as well as branch flaws. These discoveries were combined to demonstrate crack branch control through a quasi-isotropic laminate. The results have application to design of future high toughness and damage tolerant aerospace composites

Perturbation-based imperfection analysis for composite cylindrical shells buckling in compression

A numerical investigation was conducted into a perturbation-based analysis approach for assessing the imperfection sensitivity of composite cylindrical shells buckling under compression loading. The Single Perturbation Load Analysis (SPLA) approach was applied, which uses a single lateral load to introduce a realistic, worst-case and stimulating imperfection pattern. Finite element analysis was conducted for cylinders of both monolithic composite laminate and sandwich construction, with and without small and large cutouts. It was found that using a perturbation displacement equal to the shell thickness provides a suitable technique for estimating the reduction in buckling load caused by imperfections. Predictions of buckling knockdown factors using the SPLA approach showed advantages over the use of eigenmodes as the SPLA approach provides a clear design point and does not require experimental data for calibration. The effect of small and large cutouts was analogous to the effect of small and large perturbation loads. The location of the perturbation load influenced the buckling knockdown factors for both small and large cutouts, and worst-case locations were identified for both configurations

Development of a Finite Element Analysis Methodology for the Propagation of Delaminations in Composite Structures

Analysing the collapse of skin-stiffened structures requires capturing the critical phenomenon of skin-stiffener separation, which can be considered analogous to interlaminar cracking. This paper presents the development of a numerical approach for simulating the propagation of interlaminar cracks in composite structures. A degradation methodology was applied in MSC.Marc that involved modelling the structure with shell layers connected by user-defined multiple point constraints (MPCs). User subroutines were written that apply the Virtual Crack Closure Technique (VCCT) to determine the onset of crack growth, and modify the properties of the user-defined MPCs to simulate crack propagation. Methodologies for the release of failing MPCs are presented and are discussed with reference to the VCCT assumption of self-similar crack growth. Numerical results applying the release methodologies are then compared with experimental results for a double cantilever beam specimen. Based on this comparison, recommendations for the future development of the degradation model are made, especially with reference to developing an approach for the collapse analysis of fuselage-representative structures

An experimental investigation into damage modes and scale effects in CFRP open hole tension coupons

An experimental investigation was conducted to study the effect of notch size and length scale on the damage of carbon fibre-reinforced composite specimens. Open hole tension specimens in a range of configurations were tested quasi-statically to ultimate failure. The load response, damage modes and strain field development were experimentally recorded. The results demonstrated that changing the ply thickness and specimen dimensions markedly affected the damage modes and specimen behaviour. This output provides key insights into the nature of composite behaviour, and is also critical for the development and validation of analysis methodologies capturing damage initiation and progression

Development of design guidelines for postbuckling composite aerospace structures

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Structural assessment of microvascular self-healing laminates using progressive damage finite element analysis

This paper presents a progressive damage analysis methodology to numerically analyse the effect of microvascular open channels on the structural properties of self-healing fibre-polymer laminates. The tensile and compression properties of self-healing carbon-epoxy laminates containing microvascular systems are analysed using finite element models which consider progressive in-plane ply damage and intra-ply damage (matrix and delamination cracking). The models predict with good accuracy (often within 5%) the stiffness and strength of laminates containing circular or elliptical microvascular channels of different sizes and orientations. The model calculates a progressive reduction in structural properties with increasing size of microvascular channels due to increased ply waviness, which was confirmed using experimental property data. The model also predicts the location and progression of damage under increasing tensile or compression loading to final failure. The model has application as a tool for the design of microvascular systems in self-healing composites used for structural applications