Optimization of Non-Symmetric Composite Panels Using Fast Analysis Techniques

A semi-analytical approach is presented for the optimization of laminated panels with nonsymmetric lay-ups, and with the possibility of introducing requirements on the buckling load, the postbuckling response and the eigenfrequencies. The design strategy relies on the combined use of semi-analytical techniques for the structural analysis and genetic algorithms for the optimization. The structural analysis is performed with a highly efficient code based on thin plate theory, where the problem is formulated in terms of Airy stress function and out of plane displacement, expanded using trigonometric series. The solution of two distinct eigenvalue problems is performed to determine eigenfrequencies and buckling load, while an arc-length method is adopted for the postbuckling computation. The genetic algorithm is implemented by using proper alphabet cardinalities to handle different steps for the angles of orientation, while specific mutation operators are used to guarantee good reliability of the optimization. To show the potentialities of the proposed optimization toolbox, two examples are presented regarding the design of balanced non-symmetric laminates subjected to linear and nonlinear constraints. The accuracy of the analytical predictions is demonstrated by comparison with finite element results

A Fast Procedure for the Design of Composite Stiffened Panels

This paper describes the analysis and the minimum weight optimisation of a fuselage composite stiffened panel made from carbon/epoxy material and stiffened by five omega stringers. The panel investigated inside the European project MAAXIMUS is studied using a fast tool, which relies on a semi-analytical procedure for the analysis and on genetic algorithms for the optimisation. The semi-analytical approach is used to compute the buckling load and to study the post-buckling response. Different design variables are considered during the optimisation, such as the stacking sequences of the skin and the stiffener, the geometry and the cross-section of the stiffener. The comparison between finite element and fast tool results reveals the ability of the formulation to predict the buckling load and the post-buckling response of the panel. The reduced CPU time necessary for the analysis and the optimisation makes the procedure an attractive strategy to improve the effectiveness of the preliminary design phases

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

Fatigue analysis of a post-buckled composite single-stringer specimen taking into account the local stress ratio

The fatigue life prediction of post-buckled composite structures represents still an unresolved issue due to the complexity of the phenomenon and the high costs of experimental testing. In this paper, a novel numerical approach, called “Min-Max Load Approach”, is used to analyze the behavior of a composite single-stringer specimen with an initial skin-stringer delamination subjected to post-buckling fatigue compressive load. The proposed approach, based on cohesive zone model technique, is able to evaluate the local stress ratio during the delamination growth, performing, in a single Finite Element analysis, the simulation of the structure at the maximum and minimum load of the fatigue cycle. The knowledge of the actual value of the local stress ratio is crucial to correctly calculate the crack growth rate. At first, the specimen is analyzed under quasi-static loading conditions, then the fatigue simulation is performed. The results of the numerical analysis are compared with the data of an experimental campaign previously conducted, showing the capabilities of the proposed approach

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

A study on thermal buckling and mode jumping of metallic and composite plates

Composite plates in post-buckling regime can experience mode jumping in their buckling shape, suddenly increasing the number of half-waves. This phenomenon can be advantageous, be-cause the shape change could be used for local morphing or structural adaptability in future aerospace structures. A study of this phenomenon under heating is here presented, combining numerical and experimental techniques. At first, a set of parametric analysis was conducted to identify composite panels that present a mode jump when heated. Three plates were selected, one in alumi-num alloy 2024T3, and two in AS4/8552 composite material, with layup [30/−30/5/−5]s and [35/−35/10/−10]s. The plates were tested in a new test setup for thermal buckling based on low thermal expansion fixtures. The mode jumping was successfully obtained experimentally for both composite plates. Numerical simulations predicted the general trends for all plates, and the mode jumps for the composite plates

Cohesive analysis of a 3D benchmark for delamination growth under quasi-static and fatigue loading conditions

This paper evaluates the capabilities of the recently developed CF20 cohesive fatigue model, which can predict crack initiation as well as the rates of crack propagation by relying on intrinsic relationships between a stress-life diagram and its corresponding Paris law. The model is validated here using a partially reinforced double cantilever beam (R-DCB) benchmark proposed in literature. The two parameters needed for the CF20 cohesive fatigue model were obtained by performing preliminary analyses of a conventional DCB. The analysis results indicate that the CF20 cohesive fatigue model can accurately reproduce the complex evolution of the delamination observed in the R-DCB

Influence of interface ply orientation on delamination growth in composite laminates

The standard experimental procedures for determining the interlaminar fracture toughness are designed for delamination propagation in unidirectional specimens. However, in aerospace structural components, delamination usually occurs between plies at different orientations resulting in different damage mechanisms which can increase the value of the fracture toughness as the delamination propagates. Generally, numerical analyses employ the value measured at the delamination onset, leading to conservative results since the increase resistance of the delamination is neglected. In this paper, the fracture toughness and the R-curves of carbon/epoxy IM7/8552 are experimentally evaluated in coupons with delamination positioned at 0°/0° and 45°/−45° interfaces using Double Cantilever Beam (DCB) and Mixed-Mode Bending (MMB) tests. A simplified numerical approach based on the Virtual Crack Closure Technique (VCCT) is developed to simulate variable fracture toughness with the delamination length within a Finite Element code using a predefined field variable. The results of the numerical analyses compared with the experimental data in terms of load-displacement curves demonstrate the effectiveness of the proposed technique in simulating the increase resistance in delamination positioned between plies at 45°/−45° interface