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Improvements in current design approaches require further studies of the damage interaction effects of composite materials subjected to repeated out-of-plane concentrated loads. To that end, a combined simulation and experimental investigation on composite laminate under repeated indentations is reported. The repeated indentations consist of seven identical peak-force indentations that are separately applied to the centre of the laminate. The results show that delaminations grow in all seven indentations, which can be interpreted as a continuous degradation of the effective delamination growth threshold with each subsequent indentation. More specifically, the second indentation effective delamination growth threshold is 62.4 MPa, which is about 19 % lower compared to the first one (77.2 MPa). Subsequently, the delamination growth threshold degraded approximately linearly with indentation. This effective delamination growth threshold reduction can be associated with the occurrence and evolution of the crack-rich zone preceding the delamination front.

The growth of cracks between plies, i.e., delamination, in continuous fibre polymer matrix composites under cyclic-fatigue loading in operational aircraft structures has always been a very important factor, which has the potential to significantly decrease the service life of such structures. Whilst current designs are based on a 'no growth' design philosophy, delamination growth can nevertheless arise in operational aircraft and compromise structural integrity. To this end, the present paper outlines experimental and data reduction procedures for continuous fibre polymer matrix composites, based on a linear elastic fracture mechanics approach, which are capable of (a) determining and computing the fatigue crack growth (FCG) rate, da/dN, curve; (b) providing two different methods for determining the mandated worst-case FCG rate curve; and (c) calculating the fatigue threshold limit, below which no significant FCG occurs. Two data reduction procedures are proposed, which are based upon the Hartman-Schijve approach and a novel simple-scaling approach. These two different methodologies provide similar worst-case curves, and both provide an upper bound for all the experimental data. The calculated FCG threshold values as determined from both methodologies are also in very good agreement. [ABSTRACT FROM AUTHOR]

In actual operation conditions, the structural parts of automotive, aerospace and other significant industrial applications undergo the multi axial fatigue loading with non-proportionality states. Fiber metal laminates (known as FML) are the best options to manufacture these structural parts. The current study examines the effects of a specific condition of multi-axial fatigue loadings (repeated tension with static shear stresses) on the behaviors of crack growth and delamination in GLARE specimens (abbreviated of glass laminates aluminum reinforced epoxy). Analytical model with idea of maximum tangential stress criterion modification has been adopted to represent the non-proportionality state of fatigue loading. Paris equation, equivalent stress intensity factor range, fiber-bridging effects and delamination shape were included in this model. The current idea was compared with a previous idea that depends on the principal stresses in representation of non-proportional multi-axial fatigue loadings. Specimens have coding GLARE_2A 2/1 were fabricated by VARTM technique (vacuum assisted resin transfer molding technique). Special treatments have been done on the aluminum surfaces before specimen fabrication to increasing delamination resistance. Novel fatigue apparatus was manufactured to guarantee apply the under study conditions for multi axial fatigue loading and to introduce experimental investigation of crack and delamination growths in GLARE specimens. The obtained results of the current idea were in a considerable agreement with that of the previous idea of the principal stresses. The maximum deviations between analytically extracted results were less than 1% for crack growth, 5.4% for bridging stress, 4.8% for crack opening and 6.3% for delamination growth. Experimentally, the measured results have been compared with analytical results for the purpose of validation. Acceptable matching between results with maximum difference of 7.65% was the important outcome of this study.

To improve current design approaches for composite structures, it is required to further investigate the damage interaction effects of composite materials under multiple out-of-plane concentrated loads. It is first necessary to comprehend the dependence of the effective delamination threshold, which describes whether a delamination grows on pre-existing delamination damage in composite laminates. A combined experimental and numerical study is presented, in which two sequential out-of-plane quasi-static indentations are applied to fully clamped composite laminated panels, with changing distances between the two indentation locations changing. The results show that the second indentation delamination is more likely to propagate, particularly in the straight-line direction from the second indentation site to the first one, which can be interpreted as a decrease in the effective delamination threshold associated with microcracks ahead of the delamination front. The relevant percentage reduction is 37% and is independent of the imposed indentation load. As a crucial take-away, designers should be mindful that the damage interaction effects could result in greater damage than the sum of the individual cases.

Predominately, the structures and components in aeronautic, automotive and other important implementations suffer from non-proportional multi-axial fatigue loading. One of the best selected materials to fabricate these structures and components are fiber metal laminates (FML). In this research, fatigue crack-propagation in glass-laminate aluminum reinforced-epoxy (GLARE) specimen under the effect of special case of multi-axial fatigue loading (constant shear with repeated tensile stresses) has been studied analytically, numerically and experimentally. GLARE_2 A 2/1 specimen was fabricated using vacuum assisted resin transfer molding technique after special preparation aluminum surface procedure. New multi-axial fatigue apparatus including mechanical, hydraulic and control systems was manufactured to investigate crack growth experimentally. Analytical model using Paris relation was adopted to predict crack growth through aluminum layers of GLARE specimen. The rate of crack propagation is related to the mixed-mode equivalent stress intensity factor, which is estimated by the superposition for both fiber-bridging and far-field stress intensity factors. The crack opening contour, distribution of fiber-bridging stress, delamination growth and its shape were evaluated in simultaneous manner to find the fiber-bridging stress intensity factor. ABAQUS 2021 software has been used to simulate crack and delamination growth in GLARE specimen with the aforementioned fatigue loading. The extracted results from analytical and numerical models such as crack growth, crack opening, delamination shape and crack angle have been compared with that measured experimentally. The comparison indicates that there is good agreement between the calculated and measured results (with maximum deviation of 8.74% for crack growth).

Mostly the aerospace and other applications parts during their service life are subjected to the multi axial fatigue with non-proportional loading. In this study, the fatigue crack growth in glass laminate aluminum reinforced epoxy subjected to special case of non-proportional multiaxial fatigue (cyclic tension with constant shear) has been investigated. New prediction analytical model used empirical Paris equation has been adopted. Crack-growth rate was in relation with the equivalent stress-intensity factor for mixed mode at the crack-tip. Bridging stress-intensity factor was superimposed with the far-field stress-intensity factors to evaluate the equivalent stress-intensity factor. Bridging stress distribution, crack opening contour, delamination shape and its growth have been estimated simultaneously to evaluate the bridging stress-intensity factor. The obtained results indicated that there was a considerable enhancement in the crack growth rate (193.8%), number of cycles required to full failure (192%) and other performance characteristic as compared with the monolithic aluminum. Numerical simulation by ABAQUS 2021 software has been implemented to verification of the analytical model. A significant convergence in behavior between the extracted results from both analytical model and numerical simulation with maximum deviation reach to 8.2%.

[Display omitted] • The influence of the pre-existing damage on delamination growth was quantified using a proposed stress-based delamination growth prediction approach. • The underlying cause of the delamination growth in all indentations is the ongoing degradation of the effective delamination growth threshold. • The effective delamination growth threshold for the second indentation is 62.4 MPa, approximately 19 % lower than the first one (77.2 MPa). • The threshold for delamination growth decreased roughly in proportion to the loading cycle following the initial indentation. Improvements in current design approaches require further studies of the damage interaction effects of composite materials subjected to repeated out-of-plane concentrated loads. To that end, a combined simulation and experimental investigation on composite laminate under repeated indentations is reported. The repeated indentations consist of seven identical peak-force indentations that are separately applied to the centre of the laminate. The results show that delaminations grow in all seven indentations, which can be interpreted as a continuous degradation of the effective delamination growth threshold with each subsequent indentation. More specifically, the second indentation effective delamination growth threshold is 62.4 MPa, which is about 19 % lower compared to the first one (77.2 MPa). Subsequently, the delamination growth threshold degraded approximately linearly with indentation. This effective delamination growth threshold reduction can be associated with the occurrence and evolution of the crack-rich zone preceding the delamination front. [ABSTRACT FROM AUTHOR]

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In this study, the key parameters of cohesive zone modeling, which significantly affect the accuracy in delamination problems, are examined. The study is performed on various standard fracture toughness specimens subjected to mode I, mode II and mixed mode loading. The compressive buckling of a composite laminate is then modeled and the results are validated using virtual crack closure technique. It is shown that the bi-linear interface law with identified parameters provides the best match for the standard fracture toughness specimens while for buckling delamination the effect of interface law shape and interfacial strength is substantial and cannot be ignored. [ABSTRACT FROM AUTHOR]

Abstract In this study, we propose analytical and experimental methods to predict and detect deformation and interlaminar crack propagation in carbon fiber-reinforced plastics (CFRPs) based on electrical resistance measurement. The electromechanical behavior of CFRP was investigated from the elastic region to crack propagation. The CFRPs were subjected to tensile, three-point bending, and Mode I dual cantilever beam loads, and the signature electromechanical response was correlated to subsequent deformations and crack propagation. The self-sensing investigation was extended to different fiber combinations and electrode placement schemes to obtain customized sensitivity. The experimental results were verified through finite element analysis. In parallel, equivalent electrical circuit modelling was conducted to obtain the "resistor components" that exist in a CFRP and predict the electromechanical behavior under various mechanical loads. [ABSTRACT FROM AUTHOR]

Abstract Computing mode-decomposed energy release rates in arbitrarily shaped delaminations involving large fracture process zones has not been previously investigated. The J -integral is a suitable method for calculating this, because its domain-independence can be employed to reduce the integration domain to a cohesive interface, and reduce it to a line integral. However, the existing formulations for the evaluation of the mode-decomposed J -integrals rely on the assumption of negligible fracture process zones. In this work, a method for the computation of the mode-decomposed J -integrals in three-dimensional problems involving large fracture process zones and using the cohesive zone model approach is presented. The formulation is applicable to curved fronts with non-planar crack faces. A growth driving direction criterion, which takes into account the loading state at each point, is used to render the integration paths and to decompose the J -integral into loading modes. This results in curved and non-planar integration paths crossing the cohesive zone. Furthermore, its implementation into the finite element framework is also addressed. The formulation is validated against virtual crack closure technique (VCCT) and linear elastic fracture mechanics (LEFM)-based analytical solutions and the significance and generality of the formulation are demonstrated with crack propagation in a three-dimensional composite structure. Highlights • 3D J -integral applicable to problems involving large fracture process zones. • A mode I–II–III decomposed J -integral for large fracture process zones. • Evaluation of the J -integral using the information from the cohesive zone model. • Efficient implementation of the mode-decomposed J -integral using cohesive elements. • Application of the method to a 3D structure under mode I, II and III loading. [ABSTRACT FROM AUTHOR]

The delamination growth may occur in piezoelectric laminated shell with delamination under a combined electro-thermo-mechanical loading. Based on the piezoelectric theory, the variational principle of moving boundary and considering the contact effect between delamination regions, the nonlinear governing equations for the piezoelectric laminated shell with delamination under hygrothermal condition are derived, and the corresponding boundary and matching conditions are given as well. At the same time, according to the Griffith criterion, the formula of energy release rate along the delamination front is obtained and the delamination growth is studied. In the numerical calculation, the energy release rate and delamination growth of axisymmetrical piezoelectric cylindrical shell are analyzed, and a parametric study is conducted, showing the effects of geometric nonlinear, voltage, temperature and humidity, mechanical load, delamination length and depth, geometrical parameter and material property on energy release rate.

The delamination growth may occur in delaminated piezoelectric shell subjected to external dynamic load and it will further cause structural failure. Based on the variational principle of moving boundary and considering the contact effect between delamination regions, in this paper, the nonlinear governing equations for the delaminated piezoelectric shell under hygrothermal condition are derived, and the corresponding boundary and matching conditions are given. At the same time, according to the Griffith criterion, the formulas of energy release rate along the delamination front are obtained and the delamination growth is studied. In the numerical calculation, the energy release rate and delamination growth of axisymmetrical piezoelectric cylindrical shell are analyzed, and the effects of geometric nonlinear, voltage, temperature and humidity, mechanical load, delamination length and depth, geometrical parameter and boundary condition on delamination growth are discussed.

Abstract The purpose of this paper was to investigate the failure behaviors of composite stiffened panels with skin low-velocity impact (LVI) damage under axial compression. Barely visible impact damage (BVID) was introduced into the L-shaped and T-shaped stiffened panels within the skin between stiffeners. Compression after impact (CAI) tests were performed and delamination growth of specimens was observed directly. The results show that there are two types of delamination growth processes, which are unstable and stable respectively. Due to the restriction effect of the stiffener, the delamination growth is easier to initiate in the outer sublaminates on the smooth side of the skin compared to the ones on the stiffener side. In addition, according to the thermal-deply result, adjacent impact delaminations can be connected by matrix cracks, which results in a larger area of the sublaminate. It is revealed that buckling of the sublaminates and subsequent transverse delamination propagation in large area are the triggers leading to the final failure of the stiffened panel. [ABSTRACT FROM AUTHOR]

This article provides a systematic approach to accurately predict the effects of delamination deficiencies on compressive properties of composite grid-stiffened structures. The deformation of the stiffener and the skin is investigated during a mechanical test. Cohesive elements in the finite element method were employed to calculate the ultimate load-carrying capacity. The delamination deficiencies in the interface between the skin and the stiffener have a significant effect upon its ultimate load-carrying capability compared with the delamination deficiencies in the skin and in the stiffener. [ABSTRACT FROM AUTHOR]

Abstract The identification of the delamination propagation direction in three-dimensional structures with arbitrarily shaped fronts is needed in many applications. In the cohesive element framework, the propagation direction may be computed as the normal direction to a numerical damage isoline. The damage isoline tracking requires to exchange information between neighboring elements, thus post-processing global data, which is computationally expensive. This work presents a novel approach for the evaluation of the growth driving direction, only using local element information. The method can be directly implemented in a user-defined element subroutine and be evaluated at the execution time of the analysis. The presented formulation and its implementation in the commercial Finite Element code Abaqus is validated by comparison to the damage isoline shape rendering using global information. Highlights • The concept of growth driving direction for cohesive zone models is introduced. • The growth driving direction is normal to the damage isolines. • A point-wise criterion to predict the growth driving direction is formulated. • Three different alternatives of the criterion are presented and discussed. • The developed criteria can be applied to general 3D structures. [ABSTRACT FROM AUTHOR]

The introduction, originally in 2009, by the FAA of a ‘slow growth’ approach to the certification of polymer-matrix fibre composites has focused attention on the experimental data and the analytical tools needed to assess the growth of delaminations under cyclic-fatigue loads. Of direct relevance is the fact that fatigue tests on aircraft composite components and structures reveal that no, or only little, retardation of the fatigue crack growth (FCG) rate occurs as delamination/impact damage grows. Therefore, of course, the FCG data that are ascertained in laboratory tests, and then employed as a material-allowable property to design and life the structure, as well as for the development, characterisation and comparison of composite materials, must also exhibit no, or only minimal, retardation. Now, in laboratory tests the double-cantilever beam (DCB) test, using a typical carbon-fibre reinforced-plastic (CFRP) aerospace composite, is usually employed to obtain fracture-mechanics data under cyclic-fatigue Mode I loading. However, it is extremely difficult to perform such DCB fatigue tests without extensive fibre-bridging developing across the crack faces. This fibre-bridging leads to significant retardation of the FCG rate. Such fibre-bridging, and hence retardation of the FCG, is seen to arise even for the smallest values of the pre-crack extension length, a p  −  a 0 , that are typically employed. The results from the DCB tests also invariably exhibit a relatively large degree of inherent scatter. Thus, a methodology is proposed for predicting an ‘upper-bound’ FCG curve from the laboratory test data which is representative of a composite laminate exhibiting no, or only very little, retardation of the FCG rate under fatigue loading and which takes into account the inherent scatter. To achieve this we have employed a novel methodology, based on using a variant of the Hartman-Schijve equation, to access this ‘upper-bound’ FCG rate curve, which may be thought of as a material-allowable property and which is obtained using an ‘A basis’ statistical approach. Therefore, a conservative ‘upper-bound’ FCG curve may now be calculated from the DCB laboratory test data for material development, characterisation and comparative studies, and for design and lifing studies. [ABSTRACT FROM AUTHOR]

This study measured the compression after impact strength of IM7 carbon fiber laminates made from epoxy resins with various mode I and mode II toughness values to observe the effects of these toughness values on the resistance to damage formation and subsequent residual compression strength-carrying capabilities. A total of seven different epoxy resin systems were used ranging in approximate GI values of 245–665 J/m2 and approximate GII values of 840–2275 J/m2. The results for resistance to impact damage formation showed that there was a direct correlation between GII and the planar size of damage as measured by thermography. Subsequent residual compression strength testing suggested that GI had no influence on the measured values and most of the difference in compression strength was directly related to the size of damage. Thus, delamination growth assumed as an opening type of failure mechanism does not appear to be responsible for the loss of compression strength. [ABSTRACT FROM AUTHOR]

The growth of delaminations in polymer-matrix fibre composites under cyclic-fatigue loading in operational aircraft structures has always been a very important factor which has the potential to significantly affect the service-life of such structures. The recent introduction by the Federal Aviation Administration (FAA) of a ‘slow growth’ approach to the certification of composites has further focused attention on the experimental data and the analytical tools needed to assess the growth of delaminations under fatigue loads. Specific attention is given to the test and data-reduction procedures required to determine a ‘valid’ rate of fatigue crack growth (FCG), da/dN, versus the range of the energy release-rate, ΔG, (or the maximum energy release-rate, G max , in a cycle) relationship (a) to characterise and compare different types of composites, and (b) for designing and lifing in-service composite structures. Now, fibre-bridging may occur behind the tip of the advancing delamination and may cause very significant retardation of the FCG rate. Such retardation effects cannot usually be avoided when using the Mode I double-cantilever beam test to ascertain experimentally the fatigue behaviour of composites, so that a means of estimating a valid (i.e. ideally a ‘retardation-free’ or, at least, a very low-retardation) relationship is needed. The present paper presents a novel methodology, that is based on a variant of the Hartman-Schijve equation, to ascertain a valid, ‘retardation-free’, upper-bound FCG rate curves. [ABSTRACT FROM AUTHOR]

An analytical framework which incorporates damage propagation/growth into the general structural stability analysis is presented. Therefore, the conventional total potential energy approach is extended by introducing an extended total potential energy-like functional capable of describing inelastic processes in which equilibrium holds between available and the required force for producing a change in structure. The work deals with systems which are described by I generalized coordinates and K damage parameters. The damage parameters are found to be functions of I generalized coordinates and M load parameters. The underlying variational principle for inelastic solids may be solved using discrete formulations or approximate methods such as a Rayleigh–Ritz formulation. This leads to a set of non-linear algebraic equations, comprising post-critical equilibrium paths and damage propagation. In order to verify the framework, it is applied to the well-known problem in which a delaminated composite strut/plate is subjected to an in-plane compressive load. [ABSTRACT FROM AUTHOR]

Whilst composite materials are now widely used in aircraft structural components for certification purposes current designs are such that any delamination/disbond will not grow. Despite this limit fleet data and data obtained from full scale fatigue tests reveal that small sub mm initial delaminations/disbonds can grow when subjected to operational flight loads. To account for this the US Federal Aviation Authority advisory circular AC20-107B outlines a slow growth approach for certifying composite/bonded aircraft structures. A key point in AC20-107B is that growth must be both slow and predictable. Since the life of an airframe is determined by the growth of the fastest (lead) crack/delamination/disbond any analysis tool developed for a composite/bonded airframe requires the ability to predict the both the scatter in delamination/disbond growth and thereby the growth of the lead delamination/disbond in the airframe. To this end the present paper reveals that the Hartman-Schijve variant of the NASGRO equation would appear to show promise for capturing the scatter seen in Mode I delamination tests associated with specimens fabricated from nominally identical material supplied by two different suppliers. It is shown that the values of the constants in the Hartman-Schijve equation associated with these two sources are consistent and the scatter in growth can captured by allowing for small variations in the fatigue threshold term. [ABSTRACT FROM AUTHOR]

In this paper, a novel numerical procedure for delamination growth simulation able to evaluate the Fibre Bridging effect has been used to numerically investigate the mechanical behaviour of stiffened composite panels. The proposed methodology, implemented in ANSYS® FEM code by Ansys Parametric Design Language (APDL), allows to take into account the real variation of the Critical Energy Release Rate associated to the Fibre Bridging phenomenon. A first numerical application has been performed for preliminary validation purposes, simulating the flexural behaviour of representative section of a stiffened composite plate. The considerable agreement between the obtained numerical results and the literature experimental data confirms the effectiveness and the accuracy of the proposed model. Finally, a parametric analysis has been carried out in order to study the influence of the geometrical parameters of single stiffened composite panels in terms of skin-stringer debonding including Fibre Bridging. [ABSTRACT FROM AUTHOR]

Delamination of laminates around the hole edge is one of the most common but most threatening forms of damage in composites. The numerical solutions based on the finite element method (FEM) are difficult to meet the requirements of online evaluation and decision-making of the structure operation in terms of prediction efficiency. In this study, an end-to-end spatiotemporal prediction model with Encoder-Translator-Decoder architecture is developed to rapidly predict the delamination growth of laminates under compressive loading. The prediction results of single-hole delamination growth at different locations show that the model has good accuracy and generalization. The trained model can complete an inference prediction at the millisecond level, which is hard to achieve with FEM. Furtherly, a transfer framework is designed to promote the forecasting of multi-hole delamination growth. The framework is designed with a coupling coding block for learning the complex relationship of delamination damage superposition, which can effectively accelerate the training and achieve better prediction accuracy and inference speed. The developed deep learning model has good potential to be extended and applied to predict the spatiotemporal evolution of other physical fields. [Display omitted] • A spatiotemporal prediction model was developed to predict the delamination growth of laminates. • The model can be used as a reliable agent of FEM to predict delamination growth. • The model has a faster prediction speed compared to FEM, which takes only 58.1 ms. • A transfer framework was designed to apply predictions for multi-hole delamination growth. [ABSTRACT FROM AUTHOR]

[Display omitted] • The effect of the neighbouring pre-existing damage on delamination growth is quantified using a proposed stress-based delamination growth prediction method. • The stress field is determined using a finite element model, and damages are detected and quantified through ultrasonic scanning and optical microscopy. • The delamination that occurs second in the multiple-indented composite laminate is more likely to propagate at the ply interface. • The more easily propagated delamination could result from a degradation of the delamination growth threshold caused by the presence of the neighbouring damage. To improve current design approaches for composite structures, it is required to further investigate the damage interaction effects of composite materials under multiple out-of-plane concentrated loads. It is first necessary to comprehend the dependence of the effective delamination threshold, which describes whether a delamination grows on pre-existing delamination damage in composite laminates. A combined experimental and numerical study is presented, in which two sequential out-of-plane quasi-static indentations are applied to fully clamped composite laminated panels, with changing distances between the two indentation locations changing. The results show that the second indentation delamination is more likely to propagate, particularly in the straight-line direction from the second indentation site to the first one, which can be interpreted as a decrease in the effective delamination threshold associated with microcracks ahead of the delamination front. The relevant percentage reduction is 37% and is independent of the imposed indentation load. As a crucial take-away, designers should be mindful that the damage interaction effects could result in greater damage than the sum of the individual cases. [ABSTRACT FROM AUTHOR]

A layerwise theory within the framework of first-order shear deformation theory is introduced to investigate the postbuckling behavior and the delamination growth of geometrically imperfect composite plates. The Ritz method is adopted and the displacement fields are assumed in a form of polynomial series leading to a reduced computational cost. The proposed method is capable of predicting both the local buckling of the delaminated sublaminates and the global buckling of the plate. Two types of delaminations are considered: through-the-width and edge delaminations. To inhibit the penetration of delaminated domain, contact constraints are applied. The debonding of adjacent plies is modeled to illustrate the effect of delamination growth on the postbuckling response of the delaminated composite plate. A parametric study is carried out to investigate the influence of size and location of delamination as well as the amplitude of imperfection on the postbuckling behavior with concurrent propagation of delamination. A three dimensional finite element analysis is also implemented. A comparison study is conducted for numerical predictions of the developed theoretical model and implemented finite element simulation. It is found that the snap-back instability can occur during the delamination growth. Moreover, the results show that predictions based upon an analysis without consideration for the delamination growth could not be reliable. [ABSTRACT FROM AUTHOR]

Buckling behaviour and delamination growth have been investigated in Carbon Fibre-Reinforced Plastic (CFRP) laminates with artificial and impact-induced delaminations when subject to four-point bending. The energy of the impact was such that the induced damage, observed using ultrasound, did not extend across the entire width of the laminates and was barely visible on the impacted face. Stereoscopic digital image correlation was used to measure the evolution of the deformation of the laminate during bending to structural failure; and the resulting full-field displacement maps and observations of failure modes from Scanning Electron Microscopy (SEM) were used to conclude that appropriately shaped and located artificial delaminations could be employed to represent damage-induced delaminations. This enabled the development of a non-linear Finite Element Analysis (FEA) incorporating a fibre/matrix constitutive model, a modified fibre/matrix failure criterion and a delamination growth criterion to examine the interaction between the buckling behaviour and delamination growth. The predictions of the surface displacements in bending were validated, with the aid of image decomposition, using the measured data fields for a crossply laminate. The model reliably predicted the load–displacement curve and the propagation of damage in laminates with a low level impact damage, which did not extend across the width of laminates, unlike in prior reported models. [ABSTRACT FROM AUTHOR]