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This paper presents a novel method for measuring the translaminar crack resistance curve of composite laminates under Mode II shear loading. A machine learning (ML)-based approach is utilized to extract the inapparent information of the crack resistance curve from the translaminar shear strength measurements obtained from simple V-notched shear tests. The entire campaign is built on the framework of the Finite Fracture Mechanics (FFM) combined with Finite Element Method (FEM). Special emphasis is made on the nonlinear mechanical behavior of composites under shear stress since the original FFM models are designed for quasi-brittle materials. With the well-trained recurrent neural network model, the Mode II R-curve of composite laminate can be obtained with un-notched and V-notched shear strength values as inputs. Experiments were conducted on carbon fiber-reinforced composites to validate the accuracy of the R-curve obtained by the proposed approach and that by the traditional compact shear test. The successful implementation of the method suggests a more convenient and low-cost way of obtaining this important damage-related parameter for composites. © 2024

Military vehicles are often subjected to dynamic loads from mines. To protect these vehicles, steel plating underneath the vehicle is applied. However, these steel plates can be quite heavy, resulting in a slower vehicle. Composite laminates, however, are much lighter and also prove to be capable of protecting these vehicles from mines. During an explosion, the energy released from the blast will be absorbed by the composite material. This often results in the delamination of plies within the laminate. Due to the delamination, bending loads will be taken over by membrane loads. This is proven advantageous for composite materials as they are stronger in membrane loading. Unfortunately, modelling sizeable composite structures with a Direct Numerical Simulation (DNS) requires the use of a lot of elements. This, in turn, results in long computational times, particularly for non-linear analyses. Multiscale modelling is a possible solution to this problem. This study explores the method of Computational Homogenisation for delamination in composite laminates as an alternative to 3D DNS modelling. Two-dimensional Shell-Interface-Shell elements (SIFS elements) are introduced on the macroscale. These double-layered shell elements consist of two stacked Mindlin-Reissner shell elements with an interface element connecting the two shells. Each integration point of a SIFS element is linked to a mesoscopic 3D coupled Representative Volume Element (cRVE), which is also split into two shells with an interface in between. By applying linear and periodic boundary conditions that incorporate the macroscopic strains on the cRVE, mesoscopic stresses are determined, leading to macroscopic stresses and the macroscopic stiffness matrix. The proposed multiscale framework is validated by a set of load cases with different ply configurations. The results are then compared to those of a 3D DNS. The multiscale framework performs reasonably well; however, it is not without its lim, Civil Engineering

In this work, a recently proposed high-cycle fatigue cohesive zone model, which covers crack initiation and propagation with limited input parameters, is embedded in a robust and efficient numerical framework for simulating progressive failure in composite laminates under fatigue loading. The fatigue cohesive zone model is enhanced with an implicit time integration scheme of the fatigue damage variable which allows for larger cycle increments and more efficient analyses. The method is combined with an adaptive strategy for determining the cycle increment based on global convergence rates. Moreover, a consistent material tangent stiffness matrix has been derived by fully linearizing the underlying mixed-mode quasi-static model and the fatigue damage update. The enhanced fatigue cohesive zone model is used to describe matrix cracking and delamination in laminates. In order to allow for matrix cracks to initiate at arbitrary locations and to avoid complex and costly mesh generation, the phantom node version of the eXtended finite element method (XFEM) is employed. For the insertion of new crack segments, an XFEM fatigue crack insertion criterion is presented, which is consistent with the fatigue cohesive zone formulation. It is shown with numerical examples that the improved fatigue damage update enhances the accuracy, efficiency and robustness of the numerical simulations significantly. The numerical framework is applied to the simulation of progressive fatigue failure in an open-hole [±45]-laminate. It is demonstrated that the numerical model is capable of accurately and efficiently simulating the complete failure process from distributed damage to localized failure., Applied Mechanics

This study provides an investigation on mode I fatigue delamination growth (FDG) with fibre bridging at different R-ratios and temperatures in carbon-fibre reinforced polymer composites. FDG experiments were first conducted at different temperatures of R-ratios 0.1 and 0.5 via unidirectional double cantilever beam (DCB) specimens. A fatigue model, employing both the strain energy release rate (SERR) range and the maximum SERR around crack front as similitude parameter, was proposed to interpret FDG behavior. The use of this model can collapse FDG data with fibre bridging at different R-ratios into one master curve, obeying well with the similitude principles. Accordingly, it was found that FDG can accelerate with elevated temperature, but decrease at sub-zero temperature. Furthermore, there are strong correlations between the fatigue model parameters and temperature using this model in FDG interpretations. Taking these correlations into account can extend the model to accurately predict FDG behavior of other temperatures. Fractographic examinations demonstrated that temperature has effects on the FDG damage mechanisms. Both fibre/matrix interfacial debonding and matrix brittle failure were observed in FDG of −40℃. Fibre/matrix interfacial debonding becomes the dominant failure in FDG of RT and 80℃. No obvious difference on the fracture morphology was identified for FDG at different R-ratios of a given temperature., Green Open Access added to TU Delft Institutional Repository ‘You share, we take care!’ – Taverne project https://www.openaccess.nl/en/you-share-we-take-care Otherwise as indicated in the copyright section: the publisher is the copyright holder of this work and the author uses the Dutch legislation to make this work public., Structural Integrity & Composites

THICK composite materials, e.g., thickness of more than 50 mm, are increasingly being used across diverse industry sectors owing to their significant advantages of weightsavings, superiormaterial properties and load-carrying capability. These materials tend to be adopted in safety-critical applications such as large primary or secondary load-bearing structures, where mechanical failures would result in serious consequences. However, various defects and damage may occur in thick composites that endanger structural integrity and safety severely. Hence to improve the maintenance, safety and reliability of these structures, it is crucial to develop inspection methods capable of defect detection and characterisation for composite structures of significant thickness. To date, the nondestructive testing and evaluation (NDT&E) of thick composite structures still remain an urgent challenge due to their material and structural complexity, significant thicknesses, and the presence of various manufacturing and in-service defects...., Structural Integrity & Composites

The present work aims at understanding the stochastic matrix crack evolution in CFRP cross-ply laminates under tension–tension fatigue loading. An experimental campaign was carried out on twenty-three specimens at different stress levels, while two optical techniques were used for the in-situ monitoring of the accumulation of transverse matrix cracks. The results showed a significant scatter in crack evolution among specimens. This stochastic behaviour was further investigated using image analysis and numerical modelling. It was found that transverse matrix cracks can be classified into the independent and dependent cracks based on a critical crack spacing. Furthermore, the severity of interaction among cracks was quantified by introducing a dependent crack ratio. Finally, a strength-based probabilistic model was proposed to describe the scattering regime of the crack evolution. The agreement between model and test results indicates that local strength variations of 90 plies are the dominant scattering source governing the initial fatigue resistance to cracking and determining the accumulation of transverse matrix cracks among specimens. These results may provide a new insight into the stochastic nature of matrix cracking in composite laminates and aid in the design of fatigue resistance properties., Structural Integrity & Composites

The increasing trend of constructing components made of composite laminates is due to the flexibility in tailoring their properties and high strength-to-weight ratio. Nevertheless, most practical components involve cutout features for fastening and these cutouts could reduce significantly the strength of the laminate.  Due to its importance, many studies were conducted to study the effect of circular cutouts however, there is lack of information regarding the effect of various cutout shapes. Therefore, this study aims to investigate the compressive failure behavior of Kevlar Epoxy and Glass Epoxy composite laminates due to the effect of cutout shape and size with variation in fiber orientations. Finite element software, ANSYS were used to simulate the deformation and failure behavior of the laminates under compressive load. Prior to that, mesh convergence analysis and numerical validation were performed. Failure analysis was conducted for various cutout shapes (square cutout, diamond cutout, and circular) and size, based on Maximum Stress Theory. The results show that the existence of the cutouts on the composite laminates have reduced up to ten times the strength of the laminated composite plates. This information in regards to the failure behavior is important when designing components made of composite laminates under compression.

In this paper, the influence of initial geometrical imperfections on thermal stability of laminated composite plates using Layer wise plate model is presented. The mathematical model assumes layer wise variation of in-plane displacements, non-linear strain-displacement relations (in von Karman sense) and linear thermo mechanical material properties. The Koiter’s model for initial geometrical imperfection is adopted. Principle of virtual displacements (PVD) is used to derive Euler-Lagrange differential equations of linearized buckling problem. The weak form is discretized using nine-node Lagrangian isoparametric finite element. The original MATLAB program is coded for finite element solution. The effects of imperfection mode and amplitude, temperature distribution, side to thickness ratio b/h, aspect ratio b/a, boundary conditions and lamination scheme on critical buckling temperature are analysed. The accuracy of the numerical model is verified by comparison with the available results from the literature and new results are presented.

Morphogenesis of living organisms could be used as inspiration to digitally augment the reinforcement of structures. The design space in terms of fiber orientation was enlarged to facilitate more complex patches with improved performance that can be 3D printed using a ghost ply optimisation. Simulations performed on optimized open-hole tensile models show that this results in specific performance improvements up to 61% for nodal displacement optimization. As a proof of concept, nodal displacement optimized samples were printed using highly anisotropic liquid crystal polymers and showed a 40% experimental improvement in nodal displacement., Aerospace Engineering

After establishing the accuracy of the present layerwise model for linear and geometrically nonlinear bending, vibration and buckling analysis of laminated composite and sandwich plates subjected to mechanical load, as well as for thermal bending and buckling of laminated composite and sandwich plates, in this paper, free vibration of geometrically imperfect composite plate is presented. The mathematical model, based on Layerwise theory of Reddy, assumes layerwise variation of in plane displacements and constant transverse displacement through the thickness, linear strain displacement relations and linear orthotropic material properties. The Koiter’s model for initial geometric imperfection is adopted. In order to include various possible imperfection modes, such as sine type, local type and global type, an imperfection function in the form of the product of trigonometric function and hyperbolic function in (x,y) plane is used. The principle of virtual displacement (PVD) is used to derive Euler-Lagrange differential equation of motion. The weak form is discretized using nine node Lagrangian isoparametric finite element. Element stiffness matrix, initial stiffness matrix and element mass matrix are evaluated using 3x3 Gauss-Legendre integration scheme, for 2D quadratic in-plane interpolation. Consistent element mass matrix is implemented in order to preserve the total mass of the element. The original MATLAB computer program is coded for the finite element solution. The effects of imperfection amplitude on fundamental frequencies is analysed. The accuracy of the numerical model is verified by comparison with the available results from the literature. It is found that geometric imperfection greatly affect the free vibration behaviour of composite laminates and should be taken into account when designing safe and functional structures.

The version of record is available online at: http://dx.doi.org/10.1007/s00466-022-02218-2, Composite materials are gaining popularity as an alternative to classical materials in many different applications. Moreover, their design is even more flexible due to the potential of additive manufacturing. Thus, one can produce a tailored composite laminate with the optimal values of some design parameters providing the desired mechanical performance. In this context, having a parametric numerical model for the mechanical response of the composite laminate is essential to compute the optimal parameters. In the present paper, the design parameters under consideration are the angles describing the orientation of the reinforcement fibers in different layers or patches of the composite laminates. We obtain a generalized solution using Proper Generalized Decomposition (PGD) which is adopted to provide solutions with explicit parametric dependence. The Tsai-Wu failure criterion is used to estimate first ply failure. In this context, Tsai-Wu criterion is used as the objective function for the optimization of the fibre orientations in the laminate. The PGD solution provides also sensitivities for a gradient-based optimization algorithm. The potentiality and efficiency of the presented approach is demonstrated through some numerical tests., Karim El-Ghamrawy was supported by the European Education, Audiovisual and Culture Executive Agency (EACEA) under the Erasmus Mundus Joint Doctorate “Simulation in Engineering and Entrepreneurship Development (SEED)", FPA 2013-0043. Pedro Díez and Sergio Zlotnik are grateful for the financial support provided by the Spanish Ministry of Economy and Competitiveness (Grant agreement No. DPI2017-85139-C2-2-R) and by the Generalitat de Catalunya (Grant agreement No. 2017-SGR-1278, and the project H2020-RISE MATH-ROCKS GA n∘777778. This work was also partially supported by the Italian Minister of University and Research through the project “A BRIDGE TO THE FUTURE: Computational methods, innovative applications, experimental validations of new materials and technologies” (No. 2017L7X3CS) within the PRIN 2017 program., Peer Reviewed, Postprint (published version)

The present study explores the failure and surface characteristics of Glass Fiber-Reinforced Polymers (GFRP). Stepwise loading was applied in this study to understand the multi-static loading effect on the laminates before final failure. The loading was set three times to reach 10 kN with loading–unloading movement before final load until failure. The results showed that the angle of the GFRP UD laminates’ position significantly impacts the system’s failure. The results were analyzed using theoretical calculation experiment analysis, and then the failure sample was identified using ASTM D3039 standard failure. The laminates with 0° layer on edge ([0/90]S laminates) underwent preliminary failure before final failure. The mechanism of stepwise loading can be used to detect the effect of preliminary failure on the laminates. The [0/90]S laminates are subjected to stress concentration on the edge due to fiber alignment and discontinued fibers in the 0-degree direction. This fiber then fails due to debonding between the fiber and the matrix. The laminates’ strength showed that [90/0]S specimens have an average higher strength with 334.45 MPa than the [0/90]S laminates with 227.8 MPa. For surface roughness, the value of Ra increases more than six times in the 0° direction and three times in the 90° direction. Moreover, shore D hardness showed that the hardness was decreased from 85.6 SD then decreased to 70.4 SD for [0/90]S and 65.9 SD for [90/0]S. The matrix debonding, layer delamination and fiber breakage were reported as the failure mode behavior of the laminates., The present study explores the failure and surface characteristics of Glass Fiber-Reinforced Polymers (GFRP). Stepwise loading was applied in this study to understand the multi-static loading effect on the laminates before final failure. The loading was set three times to reach 10 kN with loading–unloading movement before final load until failure. The results showed that the angle of the GFRP UD laminates’ position significantly impacts the system’s failure. The results were analyzed using theoretical calculation experiment analysis, and then the failure sample was identified using ASTM D3039 standard failure. The laminates with 0° layer on edge ([0/90]S laminates) underwent preliminary failure before final failure. The mechanism of stepwise loading can be used to detect the effect of preliminary failure on the laminates. The [0/90]S laminates are subjected to stress concentration on the edge due to fiber alignment and discontinued fibers in the 0-degree direction. This fiber then fails due to debonding between the fiber and the matrix. The laminates’ strength showed that [90/0]S specimens have an average higher strength with 334.45 MPa than the [0/90]S laminates with 227.8 MPa. For surface roughness, the value of Ra increases more than six times in the 0° direction and three times in the 90° direction. Moreover, shore D hardness showed that the hardness was decreased from 85.6 SD then decreased to 70.4 SD for [0/90]S and 65.9 SD for [90/0]S. The matrix debonding, layer delamination and fiber breakage were reported as the failure mode behavior of the laminates., Deutscher Akademischer Austauschdienst (DAAD)

Morphogenesis of living organisms could be used as inspiration to digitally augment the reinforcement of structures. The design space in terms of fiber orientation was enlarged to facilitate more complex patches with improved performance that can be 3D printed using a ghost ply optimisation. Simulations performed on optimized open-hole tensile models show that this results in specific performance improvements up to 61% for nodal displacement optimization. As a proof of concept, nodal displacement optimized samples were printed using highly anisotropic liquid crystal polymers and showed a 40% experimental improvement in nodal displacement., Aerospace Engineering

After establishing the accuracy of the present layerwise model for linear and geometrically nonlinear bending, vibration and buckling analysis of laminated composite and sandwich plates subjected to mechanical load, as well as for thermal bending and buckling of laminated composite and sandwich plates, in this paper, free vibration of geometrically imperfect composite plate is presented. The mathematical model, based on Layerwise theory of Reddy, assumes layerwise variation of in plane displacements and constant transverse displacement through the thickness, linear strain displacement relations and linear orthotropic material properties. The Koiter’s model for initial geometric imperfection is adopted. In order to include various possible imperfection modes, such as sine type, local type and global type, an imperfection function in the form of the product of trigonometric function and hyperbolic function in (x,y) plane is used. The principle of virtual displacement (PVD) is used to derive Euler-Lagrange differential equation of motion. The weak form is discretized using nine node Lagrangian isoparametric finite element. Element stiffness matrix, initial stiffness matrix and element mass matrix are evaluated using 3x3 Gauss-Legendre integration scheme, for 2D quadratic in-plane interpolation. Consistent element mass matrix is implemented in order to preserve the total mass of the element. The original MATLAB computer program is coded for the finite element solution. The effects of imperfection amplitude on fundamental frequencies is analysed. The accuracy of the numerical model is verified by comparison with the available results from the literature. It is found that geometric imperfection greatly affect the free vibration behaviour of composite laminates and should be taken into account when designing safe and functional structures.

In this paper, the influence of initial geometrical imperfections on thermal stability of laminated composite plates using Layer wise plate model is presented. The mathematical model assumes layer wise variation of in-plane displacements, non-linear strain-displacement relations (in von Karman sense) and linear thermo mechanical material properties. The Koiter’s model for initial geometrical imperfection is adopted. Principle of virtual displacements (PVD) is used to derive Euler-Lagrange differential equations of linearized buckling problem. The weak form is discretized using nine-node Lagrangian isoparametric finite element. The original MATLAB program is coded for finite element solution. The effects of imperfection mode and amplitude, temperature distribution, side to thickness ratio b/h, aspect ratio b/a, boundary conditions and lamination scheme on critical buckling temperature are analysed. The accuracy of the numerical model is verified by comparison with the available results from the literature and new results are presented.

The increasing trend of constructing components made of composite laminates is due to the flexibility in tailoring their properties and high strength-to-weight ratio. Nevertheless, most practical components involve cutout features for fastening and these cutouts could reduce significantly the strength of the laminate.  Due to its importance, many studies were conducted to study the effect of circular cutouts however, there is lack of information regarding the effect of various cutout shapes. Therefore, this study aims to investigate the compressive failure behavior of Kevlar Epoxy and Glass Epoxy composite laminates due to the effect of cutout shape and size with variation in fiber orientations. Finite element software, ANSYS were used to simulate the deformation and failure behavior of the laminates under compressive load. Prior to that, mesh convergence analysis and numerical validation were performed. Failure analysis was conducted for various cutout shapes (square cutout, diamond cutout, and circular) and size, based on Maximum Stress Theory. The results show that the existence of the cutouts on the composite laminates have reduced up to ten times the strength of the laminated composite plates. This information in regards to the failure behavior is important when designing components made of composite laminates under compression.

As a primary component of the marine propulsion systems, ship propellers have been traditionally made of nickel-aluminium-bronze (NAB) or manganese bronze (MB). With the development of fibre reinforced composite materials, the advanced plastic materials are considered to be applied in the manufacturing of marine propellers. Compared to conventional rigid propellers, the composite marine propellers are expected to possess advantages of lighter weight, lower maintenance costs, higher cavitation inception speed, declined acoustic signature, and improved efficiency at off-design conditions. These potential benefits have promoted numerical and (fewer) experimental investigations of composite marine propellers. This supports the promising future of their applications, however, a more profound understanding of the underlying mechanical properties of composite marine propellers is essential before they find a wider use in practical applications in engineering., Aerospace Structures & Computational Mechanics

Interleaving composite laminates with nanofibrous mat is one of the most reliable methods for increasing interlaminar fracture toughness. The present study seeks to find out how the damage mechanisms of carbon fiber reinforced polymers (CFRPs), subjected to the mode-I and mode-II fracture tests, are affected while those are modified by interleaved Polyamide 66 (PA66) electrospun layers. For this goal, acoustic emission (AE) and scanning electron microscope (SEM) techniques were used for assessing the damage mechanisms. The mode-I test results showed that adding nanofibers could decrease matrix cracking, fiber breakage, and fiber/matrix debonding by 92%, 27%, and 87%, respectively. The AE demonstrated that no fiber breakage occurred during mode-II loading in both non-modified and nanomodified specimens which was validated by SEM images. On the other hand, the two other damage modes, i.e. matrix cracking and fiber/matrix debonding, decreased about 97% in the nanomodified laminates., Green Open Access added to TU Delft Institutional Repository ‘You share, we take care!’ – Taverne project https://www.openaccess.nl/en/you-share-we-take-care Otherwise as indicated in the copyright section: the publisher is the copyright holder of this work and the author uses the Dutch legislation to make this work public., Structural Integrity & Composites

As a primary component of the marine propulsion systems, ship propellers have been traditionally made of nickel-aluminium-bronze (NAB) or manganese bronze (MB). With the development of fibre reinforced composite materials, the advanced plastic materials are considered to be applied in the manufacturing of marine propellers. Compared to conventional rigid propellers, the composite marine propellers are expected to possess advantages of lighter weight, lower maintenance costs, higher cavitation inception speed, declined acoustic signature, and improved efficiency at off-design conditions. These potential benefits have promoted numerical and (fewer) experimental investigations of composite marine propellers. This supports the promising future of their applications, however, a more profound understanding of the underlying mechanical properties of composite marine propellers is essential before they find a wider use in practical applications in engineering., Aerospace Structures & Computational Mechanics

Interleaving composite laminates with nanofibrous mat is one of the most reliable methods for increasing interlaminar fracture toughness. The present study seeks to find out how the damage mechanisms of carbon fiber reinforced polymers (CFRPs), subjected to the mode-I and mode-II fracture tests, are affected while those are modified by interleaved Polyamide 66 (PA66) electrospun layers. For this goal, acoustic emission (AE) and scanning electron microscope (SEM) techniques were used for assessing the damage mechanisms. The mode-I test results showed that adding nanofibers could decrease matrix cracking, fiber breakage, and fiber/matrix debonding by 92%, 27%, and 87%, respectively. The AE demonstrated that no fiber breakage occurred during mode-II loading in both non-modified and nanomodified specimens which was validated by SEM images. On the other hand, the two other damage modes, i.e. matrix cracking and fiber/matrix debonding, decreased about 97% in the nanomodified laminates., Green Open Access added to TU Delft Institutional Repository ‘You share, we take care!’ – Taverne project https://www.openaccess.nl/en/you-share-we-take-care Otherwise as indicated in the copyright section: the publisher is the copyright holder of this work and the author uses the Dutch legislation to make this work public., Structural Integrity & Composites

The interlaminar fracture toughness of two unidirectional Graphite/Epoxy composite material systems has been experimentally assessed. The systems studied were prepreg composite and prepreg composite treated with carbon nanotubes (CNT). The fracture toughness has been quantified in Mode I, Mode II and Mode I/II by performing DCB, ENF and MMB tests according to relevant ISO and ASTM standards. The effect of aging by storage under hot-wet conditions has been quantified by studying these systems at room temperature without aging and at 70 °C after aging treatment. Experimental data are reported in a 3- or 5-specimen batch mode, indicating non-linear behavior and sensitivity to imperfections in coupons alignment and load application. Moreover, intermediate variables required for the estimation of fracture toughness are presented in order to be used as a reference guide for principal fracture test data evaluation. In the case of the RT systems, measured data have been compared with analytical solutions and finite element model predictions yielding good correlation for DCB and ENF tests and considerable deviation in the case of MMB tests. Main findings include that CNT-doping leads to an increase of fracture toughness in all modes, especially in Mode II, and that aging leads to less variation in measurements for both systems, indicating a more uniform matrix response., The current work has received funding from EU Horizon 2020 Clean Sky II project SHERLOC (Structural Health Monitoring, Manufacturing and Repair Technologies for Life Management of Composite Fuselage) under Grant Agreement No CS2-AIR-GAM-2014-2015-01. The authors from HAI would like to thank our ex-colleagues Dimitrios Habas and Stavros Kalogeropoulos for their major assistance with coupons fabrication and experimental work, respectively., Peer Reviewed, Postprint (author's final draft)

This work shows possibilities and limitations of the refined zigzag theory (RZT) that has been used in different structural (beam, plate and shell) finite elements. The refined zigzag theory can deal with composite laminates, adding only one nodal degree of freedom per spatial dimension of the laminate, obtaining very good accuracy. It assumes that the in-plane displacements have a piece-wise linear shape across the thickness depending on the shear stiffness of each composite layer. This paper presents the main aspects of a beam/shell of revolution element used for the numerical simulations. The details of the refined zigzag theory are given also in order to discuss some limitations that occur when dealing with the non-linear phenomenon of delamination. Two examples are presented and discussed, including different inhomogeneities that show the limitations of the RZT for the treatment of partially delaminated beams., Peer Reviewed, Postprint (author's final draft)

A progressive damage model is proposed to predict buckling strengths and failure mechanisms for both symmetric and asymmetric patch repaired carbon-fibre reinforced laminates subjected to compression without lateral restrains. Solid and cohesive elements are employed to discretize composite and adhesive layers, respectively. Coupling with three dimensional strain failure criteria, an energy-based crack band model is applied to address the softening behaviour in composites with mesh dependency elimination. Both laminar and laminate scaled failure are addressed. Patch debonding is simulated by the cohesive zone model with a trapezoidal traction–separation law applied for the ductile adhesive. Geometric imperfection is introduced into the nonlinear analysis by the first order linear buckling configuration. Regarding strengths and failure patterns, the simulation demonstrates an accurate and consistent prediction compared with experimental observations. Though shearing is the main contributor to damage initiation in adhesive, stress analysis shows that lateral deformation subsequently reverses the distribution of normal stresses which stimulates patch debonding at one of the repair sides. The influence of patch dimensions on strengths and failure mechanisms can be explained by stress distributions in adhesive and lateral deformation of repairs. Comparison between symmetric and asymmetric regarding strength and failure modes shows that structural asymmetry can intensify lateral flexibility. This resulted in earlier patch debonding and negative effects on strengths.

A progressive damage model is proposed to predict buckling strengths and failure mechanisms for both symmetric and asymmetric patch repaired carbon-fibre reinforced laminates subjected to compression without lateral restrains. Solid and cohesive elements are employed to discretize composite and adhesive layers, respectively. Coupling with three dimensional strain failure criteria, an energy-based crack band model is applied to address the softening behaviour in composites with mesh dependency elimination. Both laminar and laminate scaled failure are addressed. Patch debonding is simulated by the cohesive zone model with a trapezoidal traction–separation law applied for the ductile adhesive. Geometric imperfection is introduced into the nonlinear analysis by the first order linear buckling configuration. Regarding strengths and failure patterns, the simulation demonstrates an accurate and consistent prediction compared with experimental observations. Though shearing is the main contributor to damage initiation in adhesive, stress analysis shows that lateral deformation subsequently reverses the distribution of normal stresses which stimulates patch debonding at one of the repair sides. The influence of patch dimensions on strengths and failure mechanisms can be explained by stress distributions in adhesive and lateral deformation of repairs. Comparison between symmetric and asymmetric regarding strength and failure modes shows that structural asymmetry can intensify lateral flexibility. This resulted in earlier patch debonding and negative effects on strengths.

We numerically analyze three mechanics problems described below. For each problem, the developed computational model is verified by comparing computed results for example problems with those available in the literature. Effective utilization of single wall carbon nanotubes (SWCNTs) and single layer graphene sheets (SLGSs) as reinforcements in nanocomposites requires their strong binding with the surrounding matrix. An effective technique to enhance this binding is to functionalize SWCNTs and SLGSs by covalent attachment of appropriate chemical groups. However, this damages their pristine structures that may degrade their mechanical properties. Here, we delineate using molecular mechanics simulations effects of covalent functionalization on elastic moduli of these nanomaterials. It is found that Young's modulus and the shear modulus of an SWCNT (SLGS), respectively, decrease by about 34% (73%) and 43% (42%) when 20% (10%) of carbon atoms are functionalized for each of the four functional groups of different polarities studied. A shell theory that gives results close to the solution of the corresponding 3-dimensional problem depends upon the shell geometry, applied loads, and initial and boundary conditions. Here, by using a third order shear and normal deformable theory and the finite element method (FEM), we delineate for a doubly curved shell deformed statically with general tractions and subjected to different boundary conditions effects of geometric parameters on in-plane and transverse stretching and bending deformations. These results should help designers decide when to consider effects of these deformation modes for doubly curved shells. Composite laminates are usually fabricated by curing resin pre-impregnated fiber layers in an autoclave under prescribed temperature and pressure cycles. A challenge is to reduce residual stresses developed during this process and simultaneously minimize the cure cycle time. Here, we use the FEM and a genetic algorithm to fin

The final publication is available at Springer via http://dx.doi.org/10.1007/s12289-016-1309-4, Thermoplastic composites are widely considered in structural parts. In this paper attention is paid to squeeze flow of continuous fiber laminates. In the case of unidirectional prepregs, the ply constitutive equation is modeled as a transversally isotropic fluid, that must satisfy both the fiber inextensibility as well as the fluid incompressibility. When laminate is squeezed the flow kinematics exhibits a complex dependency along the laminate thickness requiring a detailed velocity description through the thickness. In a former work the solution making use of an in-plane-out-of-plane separated representation within the PGD – Poper Generalized Decomposition – framework was successfully accomplished when both kinematic constraints (inextensibility and incompressibility) were introduced using a penalty formulation for circumventing the LBB constraints. However, such a formulation makes difficult the calculation on fiber tractions and compression forces, the last required in rheological characterizations. In this paper the former penalty formulation is substituted by a mixed formulation that makes use of two Lagrange multipliers, while addressing the LBB stability conditions within the separated representation framework, questions never until now addressed., Peer Reviewed, Postprint (author's final draft)

This MSc thesis addresses the study of buckling and post-buckling of composite plates with elastic restraints at the edges and under any combination of in-plane loading, aiming to solve the plate response in stiffened panel structures. The implemented solutions are based on thin plate theory for mid-plane symmetric plates. The governing equations are solved using a semi-analytical formulation (not closed form) to combine advantages from analytical and numerical analysis. This approach allows to solve most of the typical laminates used in aerospace applications while allowing an improved performance when compared to FE models. The developed formulation relies on eigenbeam functions to approximate the plate behaviour for any combination of arbitrary elastic restrains, with a minimum number of degrees of freedom. This approach has proved to be able to reproduce the buckling mode and load for buckling and the out-of-plane displacement for post-buckling. The results obtained have been verified against FE commercial software package Abaqus and good to excellent agreement has been achieved using a fraction of the computational power. The relation between the ideal torsional springs and stiffeners’ restrain is approximated in order to apply the developed formulations to more practical problems involving stiffened panels. Preliminary verifications show the validity of the proposed approaches and encourages the further development of the solution to achieve a more powerful stiffened panel formulation. Moreover, the developed approaches can be extended to solve other relevant stability phenomena such as global buckling or stiffener crippling. This work is part of a Fokker Aerostructures project to develop an analytical framework for analysis and design of composite stiffened panels with post-buckling capabilities. This framework will facilitate the preliminary design of composite stiffened structures and allow further optimization without requiring the prohibitive computationa, Aerospace Engineering, Aerospace Structures and Computational Mechanics

In this paper, the performance of a recent homogenized orthotropic plasticity model for fiber reinforced composites (Vogler et al. (2013)) is investigated by comparing the model response against a micromechanical model. It is assumed that the micromechanical model which contains a recent plasticity model for polymers (Melro et al. (2013a,b)) offers a realistic representation of the plastic deformation of composite materials. Under that assumption, the performance of the homogenized model can be assessed based on the question how well it reproduces micromechanical results. Improved consistent tangent formulations for both plasticity models are presented and a study into the representative volume element size for elasto-plastic micromechanical analysis is performed. Micromechanical simulations of basic load cases are used to generate input for the homogenized model. Next, the response of the two models is compared for a range of different load cases. The loss of accuracy due to necessary simplifications in the homogenized model is quantified in the comparison. These simplifications in the mesomodel include ignoring the stress in fiber direction, disregarding the influence stress orientation under combined longitudinal shear and transverse loading, and the use of a single constant plastic Poisson ratio., Applied Mechanics

This MSc thesis addresses the study of buckling and post-buckling of composite plates with elastic restraints at the edges and under any combination of in-plane loading, aiming to solve the plate response in stiffened panel structures. The implemented solutions are based on thin plate theory for mid-plane symmetric plates. The governing equations are solved using a semi-analytical formulation (not closed form) to combine advantages from analytical and numerical analysis. This approach allows to solve most of the typical laminates used in aerospace applications while allowing an improved performance when compared to FE models. The developed formulation relies on eigenbeam functions to approximate the plate behaviour for any combination of arbitrary elastic restrains, with a minimum number of degrees of freedom. This approach has proved to be able to reproduce the buckling mode and load for buckling and the out-of-plane displacement for post-buckling. The results obtained have been verified against FE commercial software package Abaqus and good to excellent agreement has been achieved using a fraction of the computational power. The relation between the ideal torsional springs and stiffeners’ restrain is approximated in order to apply the developed formulations to more practical problems involving stiffened panels. Preliminary verifications show the validity of the proposed approaches and encourages the further development of the solution to achieve a more powerful stiffened panel formulation. Moreover, the developed approaches can be extended to solve other relevant stability phenomena such as global buckling or stiffener crippling. This work is part of a Fokker Aerostructures project to develop an analytical framework for analysis and design of composite stiffened panels with post-buckling capabilities. This framework will facilitate the preliminary design of composite stiffened structures and allow further optimization without requiring the prohibitive computationa, Aerospace Engineering, Aerospace Structures and Computational Mechanics

In this paper, the performance of a recent homogenized orthotropic plasticity model for fiber reinforced composites (Vogler et al. (2013)) is investigated by comparing the model response against a micromechanical model. It is assumed that the micromechanical model which contains a recent plasticity model for polymers (Melro et al. (2013a,b)) offers a realistic representation of the plastic deformation of composite materials. Under that assumption, the performance of the homogenized model can be assessed based on the question how well it reproduces micromechanical results. Improved consistent tangent formulations for both plasticity models are presented and a study into the representative volume element size for elasto-plastic micromechanical analysis is performed. Micromechanical simulations of basic load cases are used to generate input for the homogenized model. Next, the response of the two models is compared for a range of different load cases. The loss of accuracy due to necessary simplifications in the homogenized model is quantified in the comparison. These simplifications in the mesomodel include ignoring the stress in fiber direction, disregarding the influence stress orientation under combined longitudinal shear and transverse loading, and the use of a single constant plastic Poisson ratio., Applied Mechanics

The conceptual ideas behind isogeometric analysis (IGA) are aimed at unifying computer aided design (CAD) and finite element analysis (FEA). Isogeometric analysis employs the non-uniform rational B-spline functions (NURBS) used for the geometric description of a structure to approximate its physical response in an isoparametric sense. Due to the tensor product property of multi-variate NURBS, it is difficult to represent complex topological shapes with a single NURBS patch. Multiple, often non-conforming patches are needed to tackle increasing complexity of the geometry. To further facilitate the modeling of complex shapes and geometric features trimming technology is widely used in CAD software, however, the trimmed domain is only visually unseen and the trimming features can not be utilized directly for the analysis. To overcome these difficulties, extra efforts are needed to make isogeometric methods adapted to engineering related cases. Thin-walled structures, such as plates and shells, excel in optimal load-carrying behavior and are of major importance in the design of aerospace components and the automotive engineering. Isogeometric analysis is an ideal candidate for the modeling and simulation of shell structures, especially for rotation-free Kirchhoff-Love type shells, which profit from the exact description of the geometry and from the higher continuity properties of NURBS. Furthermore, it favorably supports continuity requirements for flexible through-the-thickness design of laminate composites. Laminated composite materials are increasingly used in the aerospace industry this asks for reliable and computationally efficient lamina theories. The classical lamination theory belongs to the class of equivalent-single-layer methods (ESL), it is computationally efficient but often fails to capture the 3D stress state accurately. The demand for an accurate 3D stress state within laminates is mainly driven by the need to identify and to evaluate potential damage o, Aerospace Structures and Computational Mechanics, Aerospace Engineering

Recent advancements in computational methods has led to improvements in the computational modeling of tensile failure of composite laminates. Models for compressive failure however, have not yet reached the same level of maturity. Particularly in the case of fiber kinking there is a need for improved modeling techniques. Fiber kinking occurs when fiber imperfection combined with matrix failure lead to localized deformation in a band of finite width (kink bands). Current computational models have focused on micro-level, where matrix and fibers are modeled independently. The various experimental and analytical research on fiber kinking has lead to numerous failure criteria and analytical models for kink bands. The collapse response of kink bands however, has yet to be incorporated in a meso-level computational model. A step is required to translate these micro-mechanical and analytical models to a meso-level framework and improve the state of the art of compressive failure modeling of composite laminates. In this thesis an attempt is made at transitioning from a micro to a meso-level failure model for fiber kinking. A discontinuous approach is proposed for modeling kink bands. The kink bands are represented as strong discontinuities in the displacement field using the phantom node method. With this method the angle of failure propagation can be easily controlled. The discontinuities are introduced after violation of a stress-based failure criterion and therefore necessitates the use of an initially rigid cohesive law to incorporate the kink band response. To construct such a law a shifted formulation is used, meaning, the law is derived with a finite initial stiffness and shifted to achieve the initially rigid behavior. Strong discontinuity analysis and a discrete micro-mechanical model are used to derive two separate cohesive models in the local kink band coordinate frame, while a simplified approach is also applied to derive a third model in the discontinuity coordi, Structural Mechanics, Structural Engineering, Civil Engineering and Geosciences

Advanced composite materials have been commonly used in aerospace engineering, because of their good mechanical properties and attractive potential for creating lightweight structures. Susceptibility to delamination is one of the most important issues in the applications of these materials. This disadvantage can prohibit the application of composite materials in primary aerospace structures and limit their lightweight potential. Therefore, characterizing fatigue delamination growth behavior in composite laminates is important for the applications of these materials in aerospace, as it provides the necessary information for the damage tolerance design philosophy. Fibre bridging is an important shielding mechanism during delamination growth in composite laminates. It can increase the fracture toughness by restraining the crack opening and inhibit delamination growth. However, there is no reliable method to take into account of its contribution to fatigue delamination growth. Thus, investigation of this phenomenon and development of a prediction method is required. The objective of this investigation is to study mode I fatigue delamination growth with fibre bridging in composite laminates and provide physics-based interpretations of fatigue delamination growth. Two approaches are applied to interpret the fatigue delamination growth behavior according to different perspectives. In the engineering perspective, the Paris relation is applied. However, in the physics-based perspective, energy principles are used. The bridging effect on fatigue delamination growth is first investigated by a series of fatigue tests. The Paris relation is used to interpret the fatigue data. It demonstrates that fibre bridging can significantly decrease the fatigue crack growth rate, making it invalid to use a single fatigue resistance curve to determine fatigue delamination with bridging. A new method, still based on the Paris relation, is developed to predict fatigue delamination growth with, Aerospace Structures and Materials, Aerospace Engineering

This paper aims to model the progressive damage of multi-bolted joints connecting structural elements made up of FRP (fiber- reinforced polymers) composite laminates and comprising different fiber materials (namely, based on basalt, carbon and glass), as well as different stacking sequences. Differences in failure mode and ultimate-load values are numerically investigated. A numerical home-made finite element model has been conceived, implemented, and validated by means of available experimental data. The numerical model is based on an incremental displacement-based approach and on a plane-stress bi-dimensional for- mulation. The stress analysis has been performed by accounting for micro-structural stress-strain localization mechanisms, and describing the progressive damage process by implementing a failure criterion operating at the constituents’ scale (namely, the Huang's criterion). Proposed results have highlighted that bolted joints based on basalt-FRP laminates and defined by a double- bolted configuration exhibited bearing failure loads comparable to those computed for glass-FRP and carbon-FRP laminates. In the case of single-bolted joints, the use of carbon-FRP laminates allowed to obtain the best mechanical properties, although joints based on basalt-FRP laminates numerically-experienced mechanical response and strength features always comparable with those of glass-FRP.

Recent advancements in computational methods has led to improvements in the computational modeling of tensile failure of composite laminates. Models for compressive failure however, have not yet reached the same level of maturity. Particularly in the case of fiber kinking there is a need for improved modeling techniques. Fiber kinking occurs when fiber imperfection combined with matrix failure lead to localized deformation in a band of finite width (kink bands). Current computational models have focused on micro-level, where matrix and fibers are modeled independently. The various experimental and analytical research on fiber kinking has lead to numerous failure criteria and analytical models for kink bands. The collapse response of kink bands however, has yet to be incorporated in a meso-level computational model. A step is required to translate these micro-mechanical and analytical models to a meso-level framework and improve the state of the art of compressive failure modeling of composite laminates. In this thesis an attempt is made at transitioning from a micro to a meso-level failure model for fiber kinking. A discontinuous approach is proposed for modeling kink bands. The kink bands are represented as strong discontinuities in the displacement field using the phantom node method. With this method the angle of failure propagation can be easily controlled. The discontinuities are introduced after violation of a stress-based failure criterion and therefore necessitates the use of an initially rigid cohesive law to incorporate the kink band response. To construct such a law a shifted formulation is used, meaning, the law is derived with a finite initial stiffness and shifted to achieve the initially rigid behavior. Strong discontinuity analysis and a discrete micro-mechanical model are used to derive two separate cohesive models in the local kink band coordinate frame, while a simplified approach is also applied to derive a third model in the discontinuity coordi, Structural Mechanics, Structural Engineering, Civil Engineering and Geosciences

This paper provides a study on mode I fatigue delamination growth in composite laminates using energy principles. Experimental data has been obtained from fatigue tests conducted on Double Cantilever Beam (DCB) specimens at various stress ratios. A concept of fatigue fracture toughness is proposed to interpret the stress ratio effect in crack growth. The fatigue fracture toughness is demonstrated to be significantly stress ratio dependent. An explanation for this phenomenon is given using SEM fractography. Fracture surface is observed to be rougher for high stress ratio in comparison with that for low stress ratio, causing the fatigue resistance increase. Therefore, the stress ratio effect in fatigue crack growth can be physically explained by a difference in resistance to crack growth., Aerospace Structures & Materials, Aerospace Engineering

Advanced composite materials have been commonly used in aerospace engineering, because of their good mechanical properties and attractive potential for creating lightweight structures. Susceptibility to delamination is one of the most important issues in the applications of these materials. This disadvantage can prohibit the application of composite materials in primary aerospace structures and limit their lightweight potential. Therefore, characterizing fatigue delamination growth behavior in composite laminates is important for the applications of these materials in aerospace, as it provides the necessary information for the damage tolerance design philosophy. Fibre bridging is an important shielding mechanism during delamination growth in composite laminates. It can increase the fracture toughness by restraining the crack opening and inhibit delamination growth. However, there is no reliable method to take into account of its contribution to fatigue delamination growth. Thus, investigation of this phenomenon and development of a prediction method is required. The objective of this investigation is to study mode I fatigue delamination growth with fibre bridging in composite laminates and provide physics-based interpretations of fatigue delamination growth. Two approaches are applied to interpret the fatigue delamination growth behavior according to different perspectives. In the engineering perspective, the Paris relation is applied. However, in the physics-based perspective, energy principles are used. The bridging effect on fatigue delamination growth is first investigated by a series of fatigue tests. The Paris relation is used to interpret the fatigue data. It demonstrates that fibre bridging can significantly decrease the fatigue crack growth rate, making it invalid to use a single fatigue resistance curve to determine fatigue delamination with bridging. A new method, still based on the Paris relation, is developed to predict fatigue delamination growth with, Aerospace Structures and Materials, Aerospace Engineering

Steel Reinforced Advanced Thin Unitized Structure (StRATUS) is a new hybrid laminate concept composed of carbon fiber composite sandwiched between thin layers of stainless steel. As a hybrid composite laminate combining layers of monolithic steel alloy and a graphite reinforced polymer matrix composite, StRATUS exhibits the potential advantages of both materials. The overall goal of this effort is to develop, analyze and demonstrate innovative structural concepts of StRATUS hybrid material to enhance the durability and damage tolerance capability of aircraft structures. Specific objectives include understanding the effects of impact damage on the hybrid composite laminates and using this knowledge to improve and optimize the StRATUS design concept., Presented at the 2014 AIAA SciTech Conference held in National Harbor, MD on 13-17 Jan 2014.

In recent years, the use of fibre-reinforced composite laminates in aerospace, mechanical, and other structural applications has become very popular due to their low density and attractive mechanical properties. However, the low inter-laminar strength of composite laminates makes them predisposed to internal damage in the form of delamination, which leads to loss of structural integrity. It is necessary to detect and assess the severity of possible delamination damage, particularly in aircraft components as early as possible in order to ensure safe operation and prevent loss of life. Detection of delaminations in composite laminates using traditional NDI techniques such as, ultrasonic testing requires aircraft to be grounded, which significantly increases maintenance cost. Implementation of structural health monitoring techniques, which can be applied online and in situ is desirable both in terms of reducing cost as well as detecting damage as soon as it occurs. Vibration monitoring offers promise in this regard.This thesis investigates the use of system identification using vibration measurements to detect and assess delaminations in composite laminates. System identification involves constructing a dynamic model of the structure through identifying system parameters from the measured dynamic response of the structure. In this study, State Space Finite Element System is employed to construct the model and the system parameters such as stiffness and damping are identified using Eigen Realization Algorithm (ERA) from the response of the structure to impulse or other forms of excitation. Damage is identified and located from unexpected variations in the system parameters, specifically, the element stiffness coefficients. A methodology is developed to assess the severity of damage by comparing the reductions in estimated stiffness values to those predicted by theory for the undamaged laminate. The proposed damage assessment technique is first applied to composite beams

In order to arrive at safe and reliable design of composite structures, understanding of the mechanisms and mechanics of damage growth in these materials is of paramount significance. Numerical models, if designed, implemented and used carefully, can be helpful not only to understand the mechanisms and mechanics of damage growth but also to predict the susceptibility of a structure to failure. In this thesis, advanced finite elements and numerical methods are explored to develop an integrated, computationally efficient and reliable numerical framework for the analysis of interacting damage mechanisms in laminated composite plates/shells subjected to transverse quasi-static and dynamic loading. A solid-like shell element is used to obtain a three-dimensional stress state in fiber-reinforced laminated composites. The element is further extended to model mesh-independent matrix cracking by incorporating a discontinuity in the shell mid-surface, shell director and thickness stretching field using the phantom node method. A progressive failure model is developed which is able to simulate impact induced damage in laminated composites. Care is taken to accurately describe the interaction between matrix cracks and delamination damage which is crucial for accurate predictions of fracture phenomena and laminate strength. Furthermore, a time-dependent progressive failure model is developed to simulate crack growth in laminated composites under dynamic loading conditions. The proposed mass discretization schemes for the solid-like shell element ensure efficient performance of the element in implicit as well as explicit elasto-fracture analysis of composite laminates. The presented numerical framework also discusses computational modeling of coupled thermo-mechanics of laminated composites in the presence of cracks. The unified computational model is able to simulate coupled adiabatic-isothermal cracks propagating arbitrarily through the finite element mesh., Structural Engineering, Civil Engineering and Geosciences

Due to the problems arising from impact damage in composite laminates, there is a need to develop fast, accurate, cost-effective and non-destructive testing methods to identify this type of damage at an early stage and thus enhance the service life of composite structures. This paper presents the results of an extensive experimental campaign conducted to investigate the feasibility of using vibration-based methods to identify damages sustained by composite laminates due to low-velocity impacts. The experimental programme included an evaluation of impact damage resistance and tolerance according to ASTM test methods, characterisation of induced damage by ultrasonic testing and quantification of the effects on the vibration response. The damage identification involved the detection, localisation, quantification and estimation of the remaining bearing capacity. Four damage indicators based on modal parameters were assessed by comparing pristine and damaged states. The results allowed for conclusions to be drawn regarding the capability and suitability of each damage indicator, including its ability to detect impact-induced damage, its precision in determining the location of damage, its sensitivity regarding damage extent and pertinent correlations with residual bearing capacity., Peer Reviewed, Postprint (published version)

This paper presents the results of an extensive experimental campaign aimed to examine the effect upon the vibration response and on the residual load-bearing capacity caused by both: isolated artificially induced interlaminar damage and low-velocity impact induced damage in composite laminates. The experimental programme included modal testing, drop-weight impact testing, ultrasonic inspection, transverse quasi-static loading testing and compression testing conducted on a set of 72 carbon fibre-reinforced composite laminated coupons. Both types of damage caused measurable changes in laminate performance, however marked divergent trends were observed. The results allowed for conclusions to be drawn regarding the adequacy of the artificial damage approach and highlighted the importance and role of other forms of degradation upon damage tolerance of laminated composites containing damage., Preprint

In order to arrive at safe and reliable design of composite structures, understanding of the mechanisms and mechanics of damage growth in these materials is of paramount significance. Numerical models, if designed, implemented and used carefully, can be helpful not only to understand the mechanisms and mechanics of damage growth but also to predict the susceptibility of a structure to failure. In this thesis, advanced finite elements and numerical methods are explored to develop an integrated, computationally efficient and reliable numerical framework for the analysis of interacting damage mechanisms in laminated composite plates/shells subjected to transverse quasi-static and dynamic loading. A solid-like shell element is used to obtain a three-dimensional stress state in fiber-reinforced laminated composites. The element is further extended to model mesh-independent matrix cracking by incorporating a discontinuity in the shell mid-surface, shell director and thickness stretching field using the phantom node method. A progressive failure model is developed which is able to simulate impact induced damage in laminated composites. Care is taken to accurately describe the interaction between matrix cracks and delamination damage which is crucial for accurate predictions of fracture phenomena and laminate strength. Furthermore, a time-dependent progressive failure model is developed to simulate crack growth in laminated composites under dynamic loading conditions. The proposed mass discretization schemes for the solid-like shell element ensure efficient performance of the element in implicit as well as explicit elasto-fracture analysis of composite laminates. The presented numerical framework also discusses computational modeling of coupled thermo-mechanics of laminated composites in the presence of cracks. The unified computational model is able to simulate coupled adiabatic-isothermal cracks propagating arbitrarily through the finite element mesh., Structural Engineering, Civil Engineering and Geosciences

In this work, the progressive failure of composite laminates is studied. Firstly, a combined elastoplastic damage model capable of representing the plastic deformations and degradation of material stiffness of composite materials has been developed based on continuum damage mechanics and plasticity theory. To simulate the strain rate-dependent effects, a consistency elasto-viscoplastic damage model that accounts for the strain rate-dependent plastic response and stiffness degradation of composites has also been developed. Based on the return mapping algorithm, implicit numerical integration procedures have been developed for the two damage models. Tangent stiffness tensors consistent with the integration algorithms have been derived separately to ensure the computational efficiency of the Newton-Raphson method for solving nonlinear problems in finite element (FE) analysis. The numerical algorithms are implemented in the FE code Abaqus through user-defined subroutines (UMATs).These damage models have been applied to the progressive failure analyses of notched and un-notched composite laminates subjected to in-plane uniaxial tensile loadings at various strain rates. It has been shown that both the combined elastoplastic damage model and the consistency elasto-viscoplastic damage model proposed in this work provide efficient tools for the progressive failure analysis of composite laminates and are capable of delivering more accurate results than several other existing models.Two individual FE models, which include the first material model for composite plies and a cohesive zone model available in Abaqus for interface layers, have been developed and implemented in progressive failure analyses of composite laminates susceptible to delaminations. The progressive failure analyses of notched laminates under in-plane tensile loading and un-notched laminates exhibiting delaminations under out-of-plane transverse impact loading have been performed. The effects of the composite

Fibre-reinforced polymer composite laminates are now commonly used in many structural applications, especially in the aerospace industry, where margins of safety are kept low in order to minimise weight. Timely detection and assessment of delaminations in composite laminates is therefore critical, as they can cause loss of structural integrity affecting the safe operation of the aircraft. Traditional NDI methods are time consuming and expensive since they require aircraft to be grounded for inspections. The current trend therefore is towards implementation of structural health monitoring systems which can monitor the structure in situ without down time. Vibration monitoring is one method which has the potential for such application. Among different methods of vibration monitoring, frequency monitoring is the simplest to implement, requiring only single point measurement, and is accurate and reliable. However, while reductions in natural frequencies clearly indicate the presence of damage, the assessment of the location and severity of the damage from measured shifts in frequencies requires the solution of the inverse problem. There have been few comprehensive studies in the past on the efficacy of delamination assessment from measured frequencies solving the inverse problem, especially for laminated plates.In this thesis, three different methods of solving the inverse problem to estimate delamination parameters from measured frequency shifts are investigated: the graphical method, artificial neural network and surrogate assisted optimisation, all applied initially to composite beams with delaminations at different interfaces. The graphical technique has been extended to the three variable problem of assessing delamination size, axial location and interface location for the first time. The artificial neural network and surrogate assisted optimisation are also applied to the five-variable problem of delaminations in plates (two dimensions and three coordinates for the

The impact damage tolerance of fiber-reinforced laminated composite materials is a source of concern, mainly due to internal induced damage which causes large reductions on the strength and stability of the structure. This paper presents a procedure based on a finite element formulation that can be used to perform numerical predictions of the impact induced internal damage in composite laminates. The procedure is based on simulating the composite performance using a micro-mechanical approach named matrix-reinforced mixing theory, a simplified version of the serial/parallel mixing theory that does not require neither the iterative procedure nor the calculation of the tangent stiffness tensor. The numerical formulation uses continuum mechanics to simulate the phenomenon of initiation and propagation of interlaminar damage with no need to formulate interface elements, resulting in a computationally less demanding formulation. To demonstrate the capability of numerical procedure when applied to a low-velocity impact problem, numerical results are compared with the experimental ones obtained in a test campaign performed on 44 laminates specimens subjected to an out-of-plane and concentrated impact event, according to ASTM test method. Results are in good agreement with experimental data in terms of delamination onset and the internal spatial distribution of induced damage., Preprint

Electrospun nanofibres as the potential reinforcement in manufacturing composite materials are recently attractive due to their simple fabrication process via electrospinning to produce continuous fibrous structures. This study concentrates on the development of novel epoxy composites laminated by layers of electrospun polylactic acid (PLA) nanofibre mats to evaluate their mechanical and thermal properties by means of flexural testing and differential scanning calorimetry (DSC), respectively. The moulded composite sheets were prepared at the fibre contents of 3 wt%, 5 wt% and 10 wt% using a solution casting method. The flexural moduli of composites have been shown to be increased by 50.8% and 24.0% for 5 wt% and 10 wt% fibre contents, respectively, as opposed to that of neat epoxy. This similar trend was also found for corresponding flexural strengths being increased by 31.6% and 4.8%. However, the flexural properties become worse at the fibre content of 3 wt% with decreases of flexural modulus by 36.9% and flexural strength by 22.9%. The examination of fractured surface morphology of composites using scanning electron microscopy (SEM) confirms a full penetration of cured epoxy matrix into electrospun PLA nanofibres despite some existences of typical fibrous structures/networks detected inside the large void cavities. On the other hand, the glass transition temperatures of composites have increased to 54-60°C due to the addition of electrospun fibres as compared to 50°C for that of epoxy, indicating that those fibrous networks may further restrict the intermolecular mobility of matrix for thermal effects.

It was shown that the damage resistance and tolerance of laminated composites can be enhanced by the employment of translaminar reinforcements (TLR) such as stitching, z-pinning and 3D weaving and also by hybrid composites. A non-dimensional analytical model focused on Mode I delamination was developed to understand the role of the TLR on delamination behavior. An explicit formula for the apparent interlaminar fracture toughness was derived in terms of the inherent fracture toughness of original materials and the bridging force due to z-pins. This model is capable of estimating the apparent fracture toughness, the bridging length and allowable bridging force thus can be useful in the design of TLR for composite laminates. Along with understanding advantages of TLR in increasing the damage tolerance of laminated composites, the damage behavior of laminated composites subjected to low velocity impact loading was studied. Based on the similarity in damage development between quasi-static and dynamic loadings observed through the short beam shear (SBS) tests, the FE analyses of the SBS specimens for quasi-static indentation and at several rates of low-velocity impact loadings were performed. The results reveal that inertia effects in the typical velocity range of the striker in SHPB, around 10m/s, can be negligible, and hence the quasi-static analysis is useful and valid in the study of damage in composite specimens under low velocity impact loading. The delamination behavior of 3D woven composites was investigated focusing on the effect of z-yarn. The 3D woven composites containing both single and double z102 yarns were chosen and compared with the 2D plain woven laminate. The double z-yarn woven composite exhibited enhanced damage tolerance compared to the single z-yarn and the plain woven laminate. The relative sliding motion between two layers is constrained by z-yarns thus the crack propagation of delamination is suppressed., The original document contains color images.