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The present study deals with the numerical modeling of the low-velocity impact damage of laminated composites which have increasingly important applications in aerospace primary structures. Such damage, generated by various sources during ground handling, substantially reduces the mechanical residual performance and the safe-service life. The purpose of this paper is to present and validate a computationally efficient approach in order to explore the effect of critical parameters on the impact damage characteristics.

Numerical modeling is considered as one of the most efficient tool as compared to the expensive and time-consuming experimental testing. In this paper, a finite element model based on explicit dynamics formulations is adopted. Hashin criterion is applied to predict the intralaminar damage initiation and evolution. The numerical analysis is performed using the ABAQUS® programme.

The employed modeling approach is validated using corresponding numerical data found in the literature and the presented results show a reasonable correlation to the available literature data. It is demonstrated that the current model can be used to capture the force-time response as well as damage parameter maps showing the intralaminar damage evolution for different impact cases with respect to the physical boundary conditions and a range of impact energies.

Low-velocity impact damage of laminated composites is still not well understood due to the complexity and non-linearity of the damage zone. The presented model is used to predict the force-time response which is considered as one of the most important parameters influencing the structural integrity. Furthermore, it is used for capturing the damage shape evolution, exhibiting a high degree of capability as a damage assessment computational tool.

Polymer composite panels are widely used in aeronautic and aerospace structures due to the high strength-to-weight ratios of these structures. The purpose of this paper is to determine the strain fields and failure mechanisms during the failure of the impacted composite laminates when subjected to compression.

A series of compression-after-impact (CAI) tests was performed on composite plates 150×100×4 mm3 made of a carbon-fibre-reinforced epoxy resin matrix. A digital image correlation and fractographic analysis by means of optical and electron microscopy are used for this purpose.

The full-field strain measurements indicate a concentrated band of compressive strain near the impact, where buckling occurs. The results indicate that the strain concentration factor can be considered to be a failure criterion. The shear strain visualisation around the impact reveals an area of heterogeneous deformation that is comparable to the detected delamination area acquired by an ultrasonic technique. Fibre and inter-fibre fractures are described for the particular impact site regions.

These experiments could improve numerical models for the CAI analyses and help to build a new criterion for this severe failure mode.

The purpose of this paper is to investigate on the behaviour of a delaminated stiffened panel; the delamination growth is simulated via fracture elements implemented in B2000++® code based on the Modified Virtual Crack Closure Technique (MVCCT), matrix cracking and fibre failure have been also taken into account.

In order to correctly apply the MVCCT on the delamination front a very fine three-dimensional (3D) mesh is required very close to the delaminated area, while a 2D-shell model has been employed for the areas of minor interest. In order to couple the shell domain to the solid one, shell-to-solid coupling elements based on kinematic constraints have been used.

Results obtained with the global/local approach are in good correlation with those obtained with experimental results.

The global/local approach based on kinematic coupling elements in conjunction with fracture elements allows to investigate and predict the behaviour of a stiffened delaminated composite panel in an efficient and effective way.

The main objective of this work is to assess the current capabilities of different commercial finite element (FE) codes in simulating the progressive damage of composite structures under quasi-static loading condition in post-buckling regime.

Progressive failure analysis (PFA) methodologies, available in the investigated FE codes, were applied to a simple test case extracted from literature consisting in a holed composite plate loaded in compression.

Results of the simulations are significantly affected by the characteristic parameters needed to feed the degradation models implemented in each code. Such parameters, which often do not have a physical meaning, have to be necessarily set upon fitting activity with an experimental database at coupon level. Concerning the test case, all the codes were found able to capture the buckling load and the failure load with a good accuracy.

This paper would to give an insight into the PFA capabilities of different FE codes, providing the guidelines for setting the degradation model parameters which are of major interest.

The purpose of this paper is to analyze the bird impact damage of fuselage composite stiffened structures by numerical method and to evaluate the damage and the bird impact resistance of different structures.

The deformation and damage of composite stiffened plates during bird impact are numerically analyzed by the explicit finite element software LS-DYNA. A comparative study on the numerical calculation results was conducted by using SPH (Smoothed Particle Hydrodynamics)-FEM (Finite Element Method) modeling and simulation. First, the I-shaped, T-shaped, straight stiffened plates and unstiffened plate were designed. Second, the accuracy of the bird model was verified and further used to evaluate bird strikes on composite stiffened plate. Third, the results of damage modes as well as displacements of the stiffened plates were compared.

The stiffeners can increase the local stiffness of the composite panel, which can effectively inhibit the bird’s movement along the impact direction. Adding stiffeners can change the panel matrix tension damage from global distribution to local distribution mode; however, the impact damage distribution and the ability to inhibit damage propagation can differ for different stiffened panels. Especially, the I-stiffened panel exhibits a better anti-bird strike performance.

The analysis of geometric parameters of structural components by numerical methods can reduce the cost of the design phase and has been widely used in aircraft design. The present study evaluated the bird impact damage of composite stiffened plates with different structures, which provides a guideline for selecting the stiffened plate structure in the fuselage skin.

This paper gives a bibliographical review of the finite element modelling and simulation of indentation testing from the theoretical as well as practical points of view. The bibliography lists references to papers, conference proceedings and theses/dissertations that were published between 1990 and 2002. At the end of this paper, 509 references are listed dealing with subjects such as, fundamental relations and modelling in indentation testing, identification of mechanical properties for specific materials, fracture mechanics problems in indentation, scaling relationship for indentation, indenter geometry and indentation testing.

The purpose of this paper is to investigate and to assess the capabilities of the most common finite element (FE)‐based tools to deal with global‐local analysis. Two kinds of coupling were investigated: shell to shell and shell to solid.

The issue of connecting non‐matching FE global and local models, characterized by different mesh refinements and/or different element types, was addressed by introducing appropriate kinematic constraints on the nodes at the interfaces. The coupling techniques available in the three FE‐based codes (ABAQUS^®, NASTRAN^® and ANSYS^®), were assessed by applying them on a common numerical test case (non‐linear buckling analysis of a square plate). Results of the global‐local simulations were compared to the results obtained for relevant reference solutions.

The continuity of displacements and stresses across the interface between global and local models and the influence of the presence of the local model on the global model solution were used as parameters to test the quality of the results. It was observed that the tools implemented in the different codes provide different results. The results characterized by a higher quality were found by using the Multi Point Constraint available in ABAQUS^®.

When dealing with complex structures, multi‐scale (global‐local) approaches are commonly adopted to optimize the computational cost by increasing mesh refinements and/or introducing elements with different formulations in specific region of the structures identified as “local model”. In this paper an overview of the coupling tools available in the main commercial FE code is given.

Examines the problems connected with quality assurance in composite material components. Focuses on the quality implications of the product during the material selection and characterization phases and on the evaluation of allowables through statistical analysis. Defines the relationship between design and the structural reliability evaluation of the loaded material and motivates the necessity of accounting for the material′s behaviour patterns with pre‐existent distributed and operation‐induced damage.

The purpose of this paper is to present microstructural and fractographic analysis of damage in aluminum (2024T3)/carbon-fiber reinforced laminates (AlC) after static tensile test. The influence of fiber orientation on the failure was studied and discussed.

The subject of examination was AlC. The fiber–metal laminates (FMLs) were manufactured by stacking alternating layers of 2024-T3 aluminum alloy (0.3 mm per sheets) and carbon/epoxy composites made of unidirectional prepreg tape HexPly system (Hexcel, USA) in [0], [± 45] and [0/90]S configuration. The fractographic analysis was carried out after static tensile test on the damage area of the specimens. The mechanical tests have been performed in accordance to ASTM D3039. The microstructural and fractographic analysis of FMLs were studied using optical (Nikon SMZ1500, Japan) and scanning electron microscope (Zeiss Ultra Plus, Germany).

FMLs based on aluminum and carbon/epoxy composite are characterized by high tensile properties depending on their individual components and the orientation of the reinforcing fibers, failure of hybrid laminates indicates the complexity process of degradation of these materials. The nature of damage in FML layers is similar to that typical in polymer composites with interlaminar delaminations, transverse cracks of the composite layers, degradation of fiber/matrix interface, damage process in FMLs is also associated mainly with interface between metal and fiber reinforced composite. The mixed damage – cohesive and adhesive – was observed.

One of the most important aspect in the designing and manufacturing process in the service life of composite structures is damage mechanisms. The damage processes in composite materials, particularly in FMLs, are more complex in comparison to metal materials and fiber reinforced polymers.

The purpose of this paper is to study the mechanical behavior in fatigue tensile mode of different cross-ply laminates constituted of unidirectional carbon fibers, hybrid fibers and glass fibers in an epoxy matrix; and to identify and characterize the local damage in the laminated materials with the use of the acoustic emission (AE) technique.

The tests in the fatigue mode permitted the determination of the effect of the stacking sequences, thickness of 90° oriented layers and reinforcement types on the fatigue mechanical behavior of the laminated materials. The damage investigation in those materials is reached with the analysis of AE signals collected from fatigue tensile tests.

The results show the effects of reinforcement type, stacking sequences and thicknesses ratio of 90° and 0° layers on the mechanical behavior. A cluster analysis of AE data is achieved and the resulting clusters are correlated with the damage mechanism of specimens under loading tests.

The analysis of AE signals collected from tensile tests of the fatigue failure load allows the damage investigation in different types of cross-ply laminates which are differentiated by the reinforcement type, stacking sequences and thicknesses ratio of 90° and 0° layers.