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

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

Assessment of Static Delamination Propagation Capabilities in Commercial Finite Element Codes Using Benchmark Analysis

With capabilities for simulating delamination growth in composite materials becoming available, the need for benchmarking and assessing these capabilities is critical. In this study, benchmark analyses were performed to assess the delamination propagation simulation capabilities of the VCCT implementations in Marc TM and MD NastranTM. Benchmark delamination growth results for Double Cantilever Beam, Single Leg Bending and End Notched Flexure specimens were generated using a numerical approach. This numerical approach was developed previously, and involves comparing results from a series of analyses at different delamination lengths to a single analysis with automatic crack propagation. Specimens were analyzed with three-dimensional and two-dimensional models, and compared with previous analyses using Abaqus . The results demonstrated that the VCCT implementation in Marc TM and MD Nastran(TradeMark) was capable of accurately replicating the benchmark delamination growth results and that the use of the numerical benchmarks offers advantages over benchmarking using experimental and analytical results

Development of a Degradation Model for the Collapse Analysis of Composite Aerospace Structures

For stiffened structures in compression the most critical damage mechanism leading to structural collapse is delamination or adhesive disbonding between the skin and stiffener. This paper presents the development of a numerical approach capable of simulating interlaminar crack growth in composite structures as a representation of this damage mecha-nism. A degradation methodology was proposed using shell layers connected at the nodes by user-defined multiple point constraints (MPCs), and then controlling the properties of these MPCs to simulate the initiation and propagation of delamination and disbonding. A fracture mechanics approach based on the Virtual Crack Closure Technique (VCCT) is used to detect growth at the delamination front. Numerical predictions using the degradation methodology were compared to experimental results for double cantilever beam (DCB) specimens to dem-onstrate the effectiveness of the current approach. Future development will focus on address-ing the apparent conservatism of the VCCT approach, and extending the application of the method to other specimen types and stiffened structures representative of composite fuselage designs. This work is part of the European Commission Project COCOMAT (Improved MA-Terial Exploitation at Safe Design of COmposite Airframe Structures by Accurate Simulation of COllapse), an ongoing four-year project that aims to exploit the large strength reserves of composite aerospace structures through more accurate prediction of collapse

Development of a Finite Element Methodology for the Collapse Analysis of Composite Aerospace Structures

In this work, an analysis methodology for capturing the critical damage mechanisms leading to collapse in composite stiffened structures is proposed. One aspect of the methodology is a global-local analysis technique that monitors a strength criterion in three-dimensional local models 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 such as a delamination or skin-stiffener debond. This approach is based on applying user-defined multi-point constraints in the skin-stiffener interface that are controlled based on the strain energy release rate as calculated using the Virtual Crack Closure Technique. A separate degradation model was also included to model the in-plane ply damage mechanisms of fibre fracture, matrix cracking and fibre-matrix shear. The complete analysis methodology was compared to experimental results for two fuselage representative composite panels tested to collapse. The two panels had different geometry and material lay-ups, where one panel was tested in an undamaged state and the other had predamage introduced as a result of cyclic loading in the postbuckling region. For both panels, the analysis methodology was shown to be capable of accurately capturing the specimen behaviour and the way in which the various damage mechanisms contributed to the final structural collapse

Development of a Finite Element Methodology for Modelling Mixed-Mode Delamination Growth

A critical failure mechanism for composite skin-stiffened structures in compression is separation of the skin and stiffener, which can be considered analogous to interlaminar cracking. This paper presents the extension of a numerical approach developed previously for simulating the propagation of interlaminar cracks in composite structures. The degradation methodology was implemented in MSC.Marc, and involves modelling the structures with shell layers connected by user-defined multi-point constraints (MPCs). User subroutines were written that apply the Virtual Crack Closure Technique to determine the onset of crack growth, and modify the properties of the user-defined MPCs to simulate crack propagation. In previous work, this model was applied only to specimens with Mode I crack growth, and two methods were proposed for handling the release of failing MPCs. In this paper, the model and release methods are extended to handle propagation in any crack growth mode. Numerical results applying the developed methodology are then compared with experimental results of fracture mechanics characterisation tests for Mode II and Mixed Mode I-Mode II. Based on this comparison, the capability of the model to represent delamination growth in any composite structure is demonstrated. Future work will focus on the application of the degradation model for the design and analysis of larger and more complex structures

Benchmark Finite Element Simulations of Postbuckling Composite Stiffened Panels

This paper outlines the CRC-ACS contribution to a software code benchmarking exercise as part of the European Commission Project COCOMAT investigating composite postbuckling stiffened panels. Analysis was carried out using MSC.Nastran (Nastran) solution sequences SOL 106 and SOL 600, Abaqus/Standard (Abaqus) and LS-Dyna, and compared to experimental data generated previously at the Technion, Israel and DLR, Germany. The finite element (FE) analyses generally gave very good comparison up to initial postbuckling, with excellent predictions of stiffness, and mostly accurate representations of the initial postbuckling mode shape, leading to fair omparison in deep postbuckling. Accurate modelling of boundary conditions and panel imperfections were crucial to achieve accurate results, with boundary conditions in particular presenting the most critical problem. Comparatively, SOL 106, SOL 600 and Abaqus gave almost identical results, whilst LS-Dyna produced less accurate results in postbuckling. The work in this paper will be compared to parallel FE analyses from other project partners, and conclusions will be made on the efficacy of various software codes for fuselage-representative composite structures

An analysis tool for design and certification of postbuckling composite aerospace structures

An experimental and numerical investigation was performed into the damage mechanisms and failure loads in skinstiffener sections. In the experimental investigation, thin strips consisting of a skin and single stiffener were cut from a range of various fuselage-representative panels. There were seven panel designs, which involved changes in geometry, lay-up, material, stiffener shape and the use of co-curing or secondary bonding to join the skin and stiffener. A total of 203 thin strip sections were cut from these panels and tested to failure. The sections were loaded in two test rigs that aimed to simulate the various symmetric and antisymmetric loads on skin-stiffener interfaces in a postbuckling panel. Five failure modes were observed, corresponding to the first damage event occurring at different locations: stiffener bend, stiffener blade, core region under the stiffener, flange edge, and skin. In general, there was good repeatability of the experimental results, particularly when classified according to failure mode, though there was a significant degree of variability in some results. For the numerical analysis, two-dimensional finite element models were analysed, and strength criteria applied in order to predict the initiation of interlaminar damage. In general, the numerical predictions gave good comparison with the experiment in terms of the critical damage locations and initiation loads, which were within the experimental scatter. Discussion is given on the sensitivity of the specimen designs, and how the twodimensional analysis approach has been applied to large fuselage-representative structures

Nonlinear static analysis of composite beams with piezoelectric actuator patches using the Refined Zigzag Theory

Piezoelectric actuators have been highly successful in a wide range of structural control applications. Assuch, there is an ongoing need for rapid and accurate structural analysis techniques, particularly for highlyheterogeneous composite materials and accounting for the actuator as a patch.Here, a new model based on the Refined Zigzag Theory (RZT) formulation that includes geometricnonlinearities is proposed for buckling, postbuckling and nonlinear static response analyses of geometricallyimperfect composite beams with piezoelectric actuators.Both the analytical and the finite element (FE) formulation are presented for symmetrically and non-symmetrically laminated beams. The FE approximation is further generalised to the case of beams withgeometric discontinuities to model composite beams with piezoelectric actuator patches. The new RZT modelis numerically verified through comparisons to Abaqus solutions for buckling and postbuckling analyses andfor the geometrically nonlinear response to an applied voltage of geometrically imperfect composite beamswith piezoelectric actuator patches.This work presents a new model for composite beams with piezoelectric actuators and confirms theremarkable advantages of RZT in terms of accuracy and computational efficiency also for challenging nonlinearanalyses, where the RZT computational time is generally less than half the time required by the FE commercial code