© 2024 International Forum on Aeroelasticity and Structural Dynamics, IFASD 2024. All rights reserved.The present work reports on the development of a numerical aerothermoelastic tool that accounts for nonlinearities of multi-physical sources to investigate the behavior of flexible wings made of a hybrid smart material. Here, hybrid materials consist of laminated composites reinforced with embedded shape memory alloy wires. The proposed model gathers geometrical, material, and aerodynamic nonlinearities to the thermal heating dynamics of SMA wires via the Joule effect. To this end, a geometrically nonlinear FE beam model is coupled with material nonlinearities via a micromechanical model that computes the homogenized properties of hybrid laminates. Nonlinear aerodynamic effects are introduced through an unsteady strip theory method in the time domain, along with the assumption of follower aerodynamic forces and a quasi-steady stall model. A set of aerothermoelastic cases was simulated by assuming various layups and SMA temperatures to tailor and analyze the aeroelastic response of hybrid wings. The outcomes have shown a considerable reduction in post-flutter oscillations as the SMA temperature increases, indicating evidence of the capability of hybrid materials for aeroelastic applications.

© 2024 International Forum on Aeroelasticity and Structural Dynamics, IFASD 2024. All rights reserved.Future aerial mobility will likely be powered by propeller propulsion, as it is more suitable for use in combination with electric motors. Therefore avoiding rotor instabilities becomes a major concern in the early project phases for the next generation of aircraft. Within this context, this work focuses on the application of Shape Memory Alloys (SMA) for Whirl Flutter (WF) suppression in propeller-driven aircraft. SMAs have a thermal-dependent modulus of elasticity, which allows the use of this class of materials to locally control the stiffness of the connections between the motor and the wing. For most of the flight, the mounting stiffness could be maintained at a minimum to better isolate the vibration coming from the motor, and only at high speeds it could be increased to avoid aeroelastic instabilities. To conduct the study, a 4 degree of freedom (dof) model of a wing section with an installed rotor was implemented and verified. This model combines a typical aeroelastic section, with springs associated with pitch and plunge dof, and the classical rotor model used in WF studies, which idealizes the rotor mounting by two torsion springs associated with pitch and yaw dof. Predictions obtained using the proposed model were compared with previous results from the literature. Following the model verification, the application of SMA was implemented by assuming that the connecting stiffness associated with the rotor installation is dependent on temperature, simulating an SMA-made mounting. Thus, it was possible to map the final flutter velocity of the system as a function of the temperatures associated with the rotor installation. The obtained results demonstrate that the flutter speed of the system may be significantly modified using this approach. They also indicate that the control of the SMA temperature shifts the dominant flutter mechanism from WF to the classical wing flutter, increasing even more the flutter speed of the system.

© 2023 Elsevier LtdAs structures become slender their non-linear aspects become more apparent and needing of assessment. In that spirit, the authors proposed a theory for addressing the effects of these non-linearities in a highly flexible beam akin to an wing in aeroservoelastic analyses regarding piezoelectric control for flutter suppression. This framework was proven quite efficient for it allowed large displacements to be naturally incorporated by means of a set of generalized variables that encoded the beam mechanics (membrane and bending) and in which space some mechanical features could be linearized. Therefore, the authors investigated the consequences of solving analytically a cantilever beam problem subjected to a material load at its free tip by means of that theory and demonstrated the connection between that problem (in particular when it comes to the buckling problem) and the Weierstrass elliptic ℘-function, a relationship not yet demonstrated to the best of the authors’ knowledge. That demonstration is the subject of this article, as well as a comprehensive study of the solutions for some loading conditions in a reference slender beam and the suggestion of further applications that could be developed from the solution found, in particular in FE analysis.

© 2023 Elsevier LtdThe aim of this study is to present a novel Semi Analytical model to analyse Mode I delamination in DCB specimens with holes, and specimens with installed fasteners. For verification and validation results were obtained with experiments and with Finite Element Analysis (FEA) based on Cohesive Zone Modelling. The proposed model obtained results with a good correlation to the experimental and FEA results, with a significant reduction in computational time. It presents the first known analytical method to include holes and fasteners in Mode I delamination analysis, and it discusses complexities and limitations of the analysis methods.

© 2023, ASM International.The present work demonstrates that a non-associated Barlat’s Yld 2000-2D plastic flow stress potential gives better correlation accuracy with the Lankford and equal biaxial coefficients of plastic anisotropy than the associated flow rule. Additionally, new generalized exact equations are presented to calculate the Lankford and equal biaxial anisotropy coefficients deduced from the Yld 2000-2D function. The investigated metals were mildly and highly anisotropic Al 2024, Al 6022, Al 2090 aluminum alloys and AISI 409 steel sheets. The non-associated Barlat’s Yld 2000-2D flow stress potential is validated by plotting on the same graph predicted r-value, normalized yield stress curves and experimental data. Newton–Raphson numerical method with a relaxation factor was employed to calculate accurately the anisotropy coefficients. Present findings for slightly and highly anisotropic aluminum alloys and AISI 409 steel revealed that Barlat’s Yld 2000-2D function can be employed for accurate characterization of metal plastic anisotropy behavior by using two independent functions: the non-associated flow stress potential and the yield stress criterion. Consequently, this procedure requires a total of 12 experimental parameters of anisotropy in calibration for accurate r-value and s-value independent curves fitting. Therefore, the proposed non-associated Barlat’s Yld 2000-12p plastic potential and yield criterion give better correlation with experimental r-value and s-value data than the associated Barlat’s Yld 2004-18p flow rule. In addition, the predicted forming limit strain curves of AISI 409 steel are in good agreement with the experimental FLC, using the non-associated Barlat’s Yld 2000-2d plastic potential, better than the associated flow potential rule.

© 2023, The Author(s), under exclusive licence to The Brazilian Society of Mechanical Sciences and Engineering.Circular Cylindrical shells structures play an important role, mainly in the aerospace sectors. In general, they are subjected to external loads and internal pressure due to internal storage such as a propulsion fuel. In this work, a semi-analytical model using Ritz method is proposed to evaluate the axial critical buckling load and internal pressure behaviour of composite cylindrical shells. Simulations were performed for different laminate stacking sequences based on unidirectional tape carbon/epoxy. Conditions of simple support and clamped edges are evaluated. The model consists of using trigonometric functions to approximate the displacement field in the Ritz formulation. In this case, the functions are chosen to meet the geometrical boundary conditions and a suitable number of terms in the Ritz method are chosen to achieve convergence results. The Ritz method formulation is based on the total potential energy and the Reissner–Mindlin hypothesis is also considered in the strain–displacement relationships for buckling cases. The critical buckling loads and buckling modes are obtained from the resultant eigenproblem when the total potential energy is minimized. The results are compared with numerical predictions obtained using the commercial software Abaqus, based on finite element method (FEM) and results available in the literature.

© 2023 Elsevier LtdFiber-reinforced polymer matrix composites are often alternative materials for aerospace structures applications where high strength and stiffness at low weight are mandatory design requirements. However, there are still open issues related to thermal effects on composite parts' mechanical properties and fracture behavior, particularly under cryogenic conditions. This work aims to investigate the fracture aspects of carbon fiber composite joints made with co-curing, co-bonding, and secondary bonding technologies when subjected to cryogenic conditions. A detailed study was carried out on their mechanical performance under Mode I and Mixed Mode I/II loading by performing interlaminar fracture tests at −54 °C. Microscopy techniques were applied to better understand the failure mechanisms observed for each bonding technology. The manufacturing process has a direct influence on the crack propagation of the laminates at cryogenic temperature. The post-cure process results in a brittle behaviour for CB and SB composite joints tested at cryogenic temperature leading to significant reductions in the fracture toughness values. The CC composite joints exhibited a tougher behaviour in comparison to CB and SB composite joints. This is mainly due to the fact that these joints are not post-cured and the presence of an interlayer that delays crack propagation.

© 2023 Elsevier LtdThis paper describes a systematic review on the propagation of translaminar damage in FRP considering different specimen configurations, data reduction schemes, fracture analysis, and mechanical properties. In particular, the influence of the specimen configuration (open-hole, edge-notched family, compact compression, compact tension, and compact tension shear) in results is highlighted. In this review, the Preferred Reporting Items for Systematic Reviews and Meta-Analyses (PRISMA) protocol was carried out as a guide. The VCCT (Virtual Crack Closure Technique) and CDM (Continuum Damage Mechanics) are the two main categories of numerical modeling approaches used to simulate the translaminar fracture in composites. Previous studies on translaminar fracture toughness characterization indicate that unwanted failure mechanisms commonly observed in opposite regions to the specimen's crack tip may be avoided by using edge-notched specimens and open-hole tensile tests combined with data reduction schemes based on Finite Fracture Mechanics criteria. Additionally, CTS (Compact Tension Shear) presents the most prominent advantage of resulting in higher stress concentration at the crack tip, avoiding unwanted compressive and other damage effects into fracture toughness behavior, adding the advantage of the mixed mode loading application and reliable R-curve. The main findings are discussed, and the shortcomings were identified to guide further investigations and provide a reference document to aid a better understanding of the benefits still to be exploited in this field.

© The Author(s) 2023.The industry of composite materials has grown in the last decade due to the requirements of light and high strength materials. The increasing demands of materials have to comply with low greenhouse gases emissions (GHG) as stated by international agreements, and reusing and recycling is a path to minimize the environmental impacts. This research aims to analyze the variables that influence the tensile strength of discontinuous laminate composites of short fibers from cutting operation process which is in the context of reusing. These variables were part of the equation of force equilibrium that involves shear strength failure criterium approach. In addition, the failure analysis and the results were compared with the literature data. Composites of glass fiber fabric wastes with varied fiber length (defined as short fiber) was designed and tested using polyethylene terephthalate (PET) as a matrix. A total of three different laminates with different fabric lengths were evaluated, totaling of seven interruptions/discontinuities along the thickness of each laminate. An image analysis of the failure sequence aided to assess the laminate behavior by comparing the stress–strain curve shape and they were in agreement with the results provided by the developed equation. The results show that this equation enables to identify the variables that influence the laminate strength: yielding stress; interface strength; stress concentration; and peel stress. In this particular research, the weak interface contributed to the low tensile strength of the laminates, and showed less influence of the “critical length,” limiting the micromechanical approach that considered a fiber filament.

© 2022 Elsevier LtdEfficient yet reliable predictive modelling tools for damage tolerance analysis became an aerospace industry requirement as composite materials provided the potential for design performance. As damage can be separated into interlaminar and intralaminar, different approaches were developed over the years. For fracture analysis of intralaminar damage, a physical link to the Linear Elastic Fracture Mechanics (LEFM) provides a reliable framework for Continuum Damage Mechanics (CDM) models. These models, however, may exhibit pathological problems related to mesh dependence, objectivity and convergence issues. Differently from Cohesive Zone Modelling (CZM) for interlaminar damage, most intralaminar models are based on phenomenological approaches for fatigue analysis and life prediction. This work provides a description of some of the current intralaminar CDM models shortcomings, related to large strain analysis and finite element topology. A novel progressive damage model is proposed where damage variables are linked to the deformation gradient. Additionally, a fatigue damage behaviour under the assumption of a Paris law for crack evolution is implemented in the proposed methodology. Both static and fatigue analysis were performed on a Compact Tension (CT) specimen geometry and compared to experimental data available on the scientific literature. Model predictions and experimental data were confronted allowing for conclusions to be drawn.

© 2022 Elsevier LtdIn recent years a considerable effort has been dedicated to the development of analysis tools that enable the design of damage tolerant structures, particularly in aerospace applications where weight reduction is a crucial requirement. One of these tools developed in recent years is the peridynamic theory, employed to solve numerically complex elastodynamics problems. One of its advantages reported in the literature is the natural ability to simulate the initiation and crack growth without the need for additional numerical procedures commonly employed in other numerical approaches, like in the conventional finite element formulation. Within this context, this paper presents a novel elastoplastic fatigue-induced damage model whose formulation is based on a strain energy framework combined with a smeared crack approach to simulate the damage process without the need of knowing the location of the crack tip and its length within the domain. The proposed model also can predict the mixed-mode damage propagation in ductile materials without knowing a priori the mixity mode ratio. This approach incorporates an analytical methodology based on the material properties that correlate the strain energy calculated away from the crack tip to the expected propagation rate predicted by the Paris Law. The accuracy of the proposed model is verified by comparing the results obtained using an in-house peridynamic FORTRAN code with the experimental results available in the open literature. Some improvements for the peridynamic material parameters are also presented in this paper, which is also verified by using the in-house code.

© 2022, The Author(s), under exclusive licence to The Brazilian Society of Mechanical Sciences and Engineering.Solid propellants are usually characterized by their ballistic and mechanical properties. However, these properties are seldom homogeneous. Processing factors such as multi-batch, casting, curing and post-curing dynamics induce transfer of loads and defects along the propellant. This propellant heterogeneity is responsible for different ballistic and mechanical properties in the grain. This paper presents a novel experimental procedure to characterize the elastic properties of single- and multi-batch solid propellants based on the use of the two-dimensional digital image correlation (DIC) method. The proposed experimental procedure has been applied to two different specimen configurations namely SBIP (single-batch inert propellant) and MBIP (multi-batch inert propellant) propellants. The SBIP specimen was manufactured in a single shot aiming at a more homogeneous mechanical behavior and uniform degree of cure along the propellant length. On the other hand, the MBIP specimen was manufactured in three different stages where each stage has a different degree of cure. Both specimens have a diameter-to-length (L/D) ratio equals to 19, which is an aspect ratio representative of typical large-scale solid-fuel grain rocket configurations. Additionally, in order to validate the in-situ measured properties, tests at small coupon level were also carried out using small cylindrical coupons taken from the same regions of interest used to measure the mechanical properties of the SBIP and MBIP specimens. A very good agreement between the measured local and global (in-situ) strain fields and mechanical properties was found in both testing scales, thus validating the proposed testing procedure based on the DIC technique. Results showed an increase in the elasticity modulus in the specimen bottom due to gravity effects.

© 2022 Elsevier LtdAdhesive bonding technologies are widely used for the assembly of stiffened panels manufactured in advanced composites for structural applications in aeronautics. However, stiffened panels are prone to the occurrence of defects or damage in the skin/stiffener junction, which will reduce the damage tolerance properties and affect structural integrity. The presence of unstable irregularities in the bonding region contributes to the decrease in the level of adhesion, limiting the resistance of the adhesive/laminate interface when subjected to mechanical loads. This article presents an experimental fracture analysis of flat panels with a longitudinal T-stiffener integrated into the skin by secondary bonding. The panels were produced in quasi-isotropic carbon/epoxy laminates with an artificial insert film replacing the adhesive film in the center of the bonding, to induce the initial damage. The tests were performed under cyclic loading followed by static axial compression loading at room temperature up to collapse. The panel selected for visual and fractographic analysis reached buckling instability with 14% of the final load, in the time interval when the failure propagation induced slight reductions in stiffness. The results obtained from this work showed the influence of the failure mechanisms combined with the formation of the failure modes and fractographic aspects that characterized the complexity of the fracture morphology provided by debonding of the skin/stiffener junction. The information revealed was relevant to the understanding of the failure process resulting from a critical defect on secondary bonding joints, applied in the integration of composite stiffened panels for aeronautics applications.

© 2022 Elsevier B.V.The vertically aligned multi-walled carbon nanotubes (VACNTs) / polydimethylsiloxane (PDMS) nanocomposite-based strain sensors presented in this study show different behavior depending on catalyst concentrations for VACNT growth. Under static tensile load, the sensor with lower catalyst concentration shows a high gauge factor (GF~1400), whereupon tunneling effect is the mechanism that dictates the sensitivity. For higher concentrations, the GF decreases (GF~40) and shows an ohmic conduction. Morphological examinations showed VACNTs are homogeneously and randomly distributed as clusters with CNT bridging in the PDMS polymer matrix. Based on dielectric impedance (DI) and direct current (DC) electrical analysis, it was possible to identify that cut-off frequency (fc) increases with VACNTs concentration. Cut-off frequency can also define high sensitivity VACNT sensors with lower VACNT density. When compared with dispersed MWCNT sensors, VACNTs have a reduced fc due to the larger internal resistance variation associated with the high tunneling effect. This effect associated with the high sensitivity of the VACNT/PDMs sensors makes it a key factor to understand the mechanisms responsible for increasing the sensitivity and manufacturing of high GF nanocomposite stretchable sensors.

© 2022 Elsevier LtdThis work describes the impact behaviour of a hybrid fibre-particulate composite composed of glass-carbon fibres and easily dispersible microparticles. The effects of fibre stacking sequence (carbon-C5, glass-G5, C2G3, G3C2, GCGCG and CG3C), particle type (silica, cement and carbon microfibers-CMF) and matrix-fibre volume fraction (40/60 and 60/40) are analysed based on a full factorial design (2 ¹4 ¹6 ¹). A drop-tower impact test characterises the hybrid composites. Fractured surfaces are examined by optical and scanning electron microscopy. The results reveal a significant synergistic effect, in which hybrid composites achieve an overall performance improvement of approximately 20% compared to glass and carbon composites. There is a greater dependence on the inclusion of particles to impact energy and resistance, reaching increased values, especially when silica particles are added. A greater amount of matrix phase ratio leads to a more efficient rheology in terms of fibre-particle interface. In addition, symmetrically placed carbon fibre layers on both sides of the beam under tensile and compressive loads (CG3C) enhance their impact performance in hybrid configuration.

© 2022 Elsevier LtdPeridynamic Theory based models allow simulating the initiation and growth of cracks in solid materials, without the aid of additional methods commonly employed in the conventional finite element formulation. Within this context, a new elastoplastic damage model is proposed to use with the Peridynamic Theory. This proposed damage model combines Von Mises plasticity-based theory with a smeared cracking approach enabling damage prediction within an energy-based framework. The formulation incorporates a mixed-mode propagation criterion to account for the effect of both axial and shear stresses in the simulation, which in turn allows prediction of damage progression in ductile materials under multiaxial loading without knowing a priori the mode mixity ratio. This proposed damage modeling approach can be used within any constitutive peridynamic model, by relying on the displacement field obtained in the simulation.

© 2022 Elsevier LtdThis paper reports the formulation of an aerothermoelastic tool developed to investigate the behavior of flexible beam-like wings made of a hybrid adaptive material. Here, hybrid materials are defined as laminated composites additionally reinforced with embedded shape memory alloy wires. As main novelties, the proposed model couples geometrical, material, and aerodynamic nonlinearities to the thermal dynamics of SMA wires undergoing Joule's effects, thereby establishing a multi-physical nonlinear problem. Geometrical nonlinearities were taken into account via an FE model of a 2D Timoshenko's beam experiencing large deformations. Material nonlinearities were incorporated by a semi-empirical micro-mechanical model that computes the properties of hybrid laminates. To complement, nonlinear aerodynamic effects were introduced via an unsteady strip theory method in the time-domain, along with a nonlinear stall model and an assumption of follower aerodynamic forces. A set of numerical aerothermoelastic cases was performed by assuming various layups and SMA temperatures, with the objective of tailoring the aeroelastic response of hybrid wings. These cases were shown to lead to a considerable reduction in both post-flutter oscillations and post-divergence amplitudes as the SMA temperature increases. The outcomes have indicated compelling evidences on the applicability of embedded SMAs for structural, shape or aeroelastic control of flexible wings.

© 2021 Elsevier LtdA semi-analytical model for buckling analysis of stiffened composite panel with debonding defect, subjected to in-plane shear load is developed and verified. The model formulation is based on the Rayleigh–Ritz method combined with the principle of total stationary potential energy. The domain is discretized ensuring the continuity C1, each domain displacement was approximated using a base of polynomial hierarchical functions. Finite element analyses and experimental tests were also performed to verify the proposed model and investigate the influence of the defect size on the panel stability. The proposed semi-analytical model is an efficient and accurate design tool that can be used in the prediction and identification of critical design scenarios for damage tolerant aerostructures.

© Proceedings of the International Forum of Aeroelasticity and Structural Dynamics 2022, IFASD 2022.This study performs an experimental and numerical investigation on the nonlinear aeroelastic response of composite flat plate-like wings with a ballast at their free tips. The effects of different chord-wise ballast positions are experimentally examined in a set of six rectangular wings laminated with different layups. The proposed numerical model brings forward a nonlinear FE beam model accounting for aerodynamic and geometrical nonlinearities. The latter were taken into account by a total Lagrangian formulation in order to describe the exact kinematics of a Timoshenko’s beam. Nonlinear aerodynamic loads were computed via an unsteady strip theory model in the time-domain, with the Jones approximation for the Wagner’s function. Additionally, a quasi-steady stall model based on an experimental quasi-static stall curve for flat plates was used to interpolate the lift-curve slope. Different nonlinear post-flutter LCO behaviors were obtained for the different ballast positions and layups tested. To conclude, the reasonable correlation between model and experiments indicated that the nonlinear approach performed here was capable to predict the aeroelastic behavior of the tested wings.

© 2022, AIAA International. All rights reserved.The effect of addition of viscoelastic plies on the acoustic scattering quadrupoles near the trailing edge of laminated plates is evaluated. A numerical method is applied to compute the acoustic field scattered by finite flexible plates. For a two-dimensional problem whereby a cantilevered plate scatters sound from a point quadrupole near the free edge, results show that adding viscoelastic layers to a composite plate can modify the far-field sound. Parametric investigations show that this treatment reduces scattered noise near resonance frequencies. Discussions on the positioning and thickness of the viscoelastic layers and operating temperature are provided. The use of outer viscoelastic layers in composite plates is predicted to significantly reduce acoustic scattering near resonances due to structural damping.

© The Author(s) 2021.In this article, the structural behavior of co-cured composite joint (CC), co-bonded composite joint (CB), and secondary-bonded composite joint (SB) under Mode II fatigue loading was evaluated. Fatigue performance was evaluated in sub-critical strain energy release rate (SERR) associated with Mode II fatigue induced delamination growth onset. Fatigue tests were carried out using the three-point bending End Notched Flexure test setup for different energy ratios. The experimental results are presented in terms of SERR versus number of cycles, and the SERR threshold for no growth is determined (Gth). Fractographic analyses were performed in order to identify the main failure mechanisms related to each joining technology under Mode II. The results indicated an initial cohesive failure followed by an adhesive failure promoted by crack propagation at the interface between the adhesive and the composite adherend on SB and CB samples, through the coalescence of microcracks that promote the adhesive failure process, leading to fiber pull-out from the matrix and cusps formation in the fracture surface. These results explain the low performance behavior observed on SB and CB bonded techniques. It is worth mentioning that the results and behavior observed in this work are valid only for the laminates, adhesives, surface treatment, and environmental conditions tested herein.

© 2021, AIAA International. All rights reserved.Recent works have addressed piezoelectrically induced stresses as a potential technique for aeroelastic stability augmentation of biclamped structures. Because active-control-based techniques are conventionally used for aeroelastic stabilization, this paper presents a comparative study on the effectiveness of piezoelectrically induced stresses and active control for aeroelastic stability augmentation. A finite-element-method-based piezoaeroelastic model is proposed employing two-node, eight-degree-of-freedom smart beam elements and an unsteady, strip-theory-based aerodynamic approach involving vertical gust components. A stability augmentation system was designed for stiffness control and used for comparison in a particular wing configuration. Analyses involving the stability margins, input signal energy, and the response to vertical gust were performed and discussed. The results showed that piezoelectrically induced stresses and active control can provide equivalent flutter speed increase for the two cases of gain margin considered. However, the system can become unstable due to control signal saturation when submitted to high-amplitude gusts, which was not observed when piezoelectrically induced stresses is used. On the other hand, it was observed that the active control required a much smaller amount of energy for stabilization. In general, it was noticed that the piezoelectrically induced stresses technique was not as effective as active control to increase the flight envelope of biclamped structures, because active control can provide equivalent aeroelastic stability improvement with smaller amount of energy. However, it is found to be a promising strategy to be used on emergency devices, where the aeroelastic stability of biclamped structures needs to be guaranteed in critical aerodynamic disturbance conditions.

© The Author(s) 2021.Carbon fiber fabrics have been largely used in composite structures as they provide high mechanical strength and potential weigh reduction, allowing more efficiency in product design. However, the production of the parts generates scraps that is discarded as a waste, becoming a challenge to recycle the carbon fiber with predictable mechanical strength. Within this context, this research analyzed strategies of laying up carbon woven fabrics based scraps, in order to reach a desirable mechanical properties in bending loading. Three types of laminates were manufactured using varied fabric size and number of discontinuities in the layup combined with polyethylene terephthalate (PET) film as a matrix. The obtained composites were tested under four-point-bending test and an energy-strength based analysis was conducted. This analysis explained a strategic position of fabric scrap to maximize the bending strength: providing a value of 106.33 MPa for a composite with high number of discontinuities against 83.11 MPa for another with less discontinuity.

© 2021 John Wiley & Sons Ltd.In this article, a nonlinear structural formulation that uses a new dimensionless set of generalized displacements is proposed for solving geometrically nonlinear beam problems, being validated by several applications given in literature where a cantilever beam is likely to undergo large displacements. By this approach, useful simplifications and insights are achievable in the analysis process, as the system matrices becoming linear and the reduction of required interpolation continuity degree. The formulation is firstly developed—in a Lagrangian perspective—and the equilibrium equations are then derived using Hamilton's principle. In the sequence, using finite element method, it is substantiated by comparison to examples given in literature in static, dynamic, and finally in an application where piezoelectric effects intervene, in order to assess its multiframework capabilities and deliver a convenient approach whereby beams constituted of smart or conventional materials can be efficiently studied.

© 2021 Elsevier LtdMode I delamination fatigue crack growth behavior was investigated in a carbon-epoxy prepreg fabric laminate by evaluating the mean load effects. The fatigue crack growth rate was determined as a function of the maximum Mode I strain energy release rate considering an exponential fitting function according to the Compliance Based Beam Method (CBBM). Classical data reduction techniques were combined with the proposed method, indicating less scatter on results and satisfying basic assumptions of smoothness and continuity for the fatigue crack growth process. The delamination growth rate curve proved to be strongly affected by the applied mean load as the fatigue onset delamination. The number of cycles to onset is higher for lower load levels and, considering the stable propagation region, a higher delamination growth rate was reported for higher load levels. The fractographic analysis has confirmed the effects of cyclic loading and the mean load levels on fracture surfaces. For a higher mean load, failure mechanisms expose static aspects and substantial presence of microcracks at fiber imprints, while, under lower mean load were noticed significant wear and plasticity and featureless fiber tracks.

© The Author(s) 2021.This paper describes a classical laminate theory-based constitutive model for portraying thermoplastic composites’ mechanical properties and the development of residual stresses during consolidation. The extended Hillier model is applied to describe the material’s crystallisation and as such is able to provide final part quality as a function of the process cooling history while taking into account the first and second crystallisation mechanisms occurring concurrently. With the developed model, a parametric study was performed taking into account layups that are commonly used in the aerospace industry, where general design guidelines are suggested. Some of the advantages of using cross-ply and quasi-isotropic laminates became clear as no shear residual stresses were predicted for those laminates. However, highly anysotropic laminates may also offer structural advantages. Numerical simulations indicate that the crystallisation residual strains can be, although smaller than thermal residual strains, relevant to final part quality. The combination of both effects may result in high residual stresses at ply level which in turn can compromise the ultimate strength of the laminates and make it difficult to attain the desired part’s geometrical tolerances.

© 2021 Elsevier LtdThis work presents an efficient explicit finite element model for predicting the nonlinear aeroelastic behavior of composite panels in the supersonic regime. The first-order shear deformation plate theory in conjunction with the von Kármán nonlinear strains is used for structural modeling and the linear piston theory is used to model the aerodynamic loads. In order to reduce the computational cost of the simulations, a lumping procedure is employed in the mass and aerodynamic damping matrices of the finite element model. No modal reduction is performed and the central difference method is used for the numerical direct integration in time of the nonlinear equations. The model is verified using results from the literature and it is demonstrated that the lumping procedure drastically reduces the computational cost of the simulations.

© 2021 Elsevier LtdThis study performs an experimental and numerical investigation on the nonlinear aeroelastic response of cantilever high-aspect-ratio beam-like wings with a ballast at their free tips, emulating the effects of a store. As an extent, the effects of different chord-wise ballast positions are experimentally examined for two highly flexible rectangular wings. Furthermore, the numerical model proposed brings forward a nonlinear finite element beam model accounting for aerodynamic nonlinearities, via stall and follower forces models, along with geometrical nonlinearities due to large displacements and rotations. A great variety of analyses were performed: First, the flutter boundaries of the wings were analyzed; second, the limit cycle oscillation amplitudes and frequencies in the oscillating wings were evaluated; third, the coupling behavior and the nonlinear responses obtained were discussed under several attributes. The geometrical nonlinearities were taken into account by a total Lagrangian formulation based on a straightforward and consistent interpolation field in order to describe the exact kinematics of a Timoshenko's beam. Nonlinear aerodynamic loads were computed via an unsteady strip theory in the time-domain with the Jones approximation for the Wagner's function along with a follower aerodynamic loads assumption. Additionally, a non-usual stall model based on an experimental quasi-static stall curve for flat plates was used to interpolate the lift-curve slope. The experimental and numerical results indicated a minimum flutter speed for ballast positions about of −5 mm toward the leading edge. Next, different nonlinear post-flutter LCO behaviors were obtained for the different ballast positions tested. To conclude, the good correlation between model and experiments indicated that the nonlinear modeling approach proposed herein was capable to predict the aeroelastic behavior of the tested high aspect-ratio wings.

© 2020 Elsevier LtdDespite composite materials presenting a viable design solution for structural weight savings, their low resistance to impact damage portraits a potential drawback. This work presents a study on the compressive pre-loading effects on Low Velocity Impact, LVI, for composite laminated plates. Despite the absence of a standard testing procedure accounting for pre-load effects, a test rig was designed specifically for LVI on pre-loaded plates while preserving ASTM D7136M-15 guidelines for the stress-free specimens. Woven fabric composite laminated plates were manufactured using the Resin Transfer Moulding (RTM) process and subjected to impact on pre-loaded conditions. Three pre-load levels were prescribed, representing conditions ranging from low strain levels until post buckling regime. FE numerical models were implemented in ABAQUS FE code, comparing different Continuum Damage Mechanics (CDM) constitutive model formulations with experimental results. Analytical and numerical predictions showed good correlation with experimental results, allowing for further application of the test apparatus and methodology.

© 2020 Elsevier LtdThis work presents an experimental study on the mechanical behaviour of thermoplastic polyarlyetherketone (PAEK) based composite laminates reinforced with woven carbon fibres, subjected to high strain rates under compression loading. The specimens were tested using a Split Hopkinson Pressure Bar (SHPB). The tests covered the working temperatures and stress envelopes of −54 °C, RT (25 °C), and 80 °C, each tested at six different off-axis angles. The high-speed imaging system was used to monitor the failure process. The strain on the loading direction was determined using Digital Image Correlation (DIC). Fractography analysis was performed to understand the influence of temperature on the damage aspects, using Scanning Electron Microscopy (SEM). The −54 °C and RT SHPB tests showed more intralaminar damage aspects than the ones tested at 80 °C, which presented a preference for the interlaminar damage aspect and exhibited the lowest strength. A new failure criterion dependent on temperature and strain rate was proposed, based on a phenomenological approach and experimental results. The fabric architecture at the ply level was idealised as a two-part mosaic model. The crystallinity of the PAEK thermoplastic matrix was analysed using Differential Scanning Calorimetry (DSC). Results showed no evidence of crystallinity degree variation induced by the high strain rate tests.

© 2021, The Minerals, Metals & Materials Society.The plastic anisotropic response of stainless steel materials is investigated in this paper by using Barlat’s Yld 2000-2d yield criterion. A new set of anisotropy coefficients is proposed and calibrated based on material experimental data. The new set of coefficients for the Lankford anisotropy coefficient, normalized yield stress, and equal biaxial stress were numerically obtained using the Newton–Raphson method. Study cases for AISI 409L and AISI 430 materials are presented and discussed. Correlations between predictions and experimental results indicate that Barlat’s yield stress criterion and plastic stress potential for stainless materials are not coincident. Hence, the Barlat’s non-associate flow rule gives better fitting with the experimental Lankford’s coefficient of anisotropy results.

© 2020 Elsevier LtdIn the open literature, many authors have used the fractional calculus in conjunction with the finite element method to model certain viscoelastic systems. The so-named fractional derivative model may be a better option for transient analyses of systems containing viscoelastic materials due to its causal behavior and its capability to fit accurately the viscoelastic damping properties and to represent properly their fading memory. However, depending on the situation, it leads to costly computations due to the integration of the non-local viscoelastic displacement and stress fields, especially for long time intervals. In this contribution, it is proposed a new and efficient general three-dimensional fractional constitutive formulation based on the use of a recurrence term to give a simplest and low-cost constitutive law to describe the frequency- and temperature-dependent behavior of viscoelastic materials, especially for complex systems. To demonstrate the efficiency and accuracy of the proposed formulation compared with those available in the literature, an academic example formed by a thin three-layer sandwich plate is performed and the main features and capabilities of the proposed methodology are highlighted.

© 2020 Society of Plastics EngineersAdditively manufactured composites have been demonstrating promising results with the development of new materials of high mechanical performance, which draws attention from several fields, for example, biomedical, electronics and aeronautics. However, as such materials are based on novel technologies, it is necessary to better understand their resulting characteristics and properties. For instance, evaluating the effect of environmental conditions on their mechanical performance is important, especially when moisture-sensitive polymers such as polyamide (PA) are employed as matrix. This work aims to understand and to characterize the moisture effect on the mechanical properties of additively manufactured Nylon and continuous carbon fiber (CF)-reinforced Nylon-based thermoplastic. Tensile and compressive tests were carried out in accordance with ASTM standards for the printed samples at their maximum moisture content and for samples submitted to drying after the saturated condition. Moreover, moisture absorption and swelling behaviors were assessed and discussed. The experimental results showed that moisture significantly affects the fiber/matrix interface, as well as the adhesion between printed filaments. These changes led to a decrease in the general mechanical properties in saturated state, including those in the fiber direction. Furthermore, a permanent degradation was observed in some properties after drying. Thus, the importance of considering water content and aging effect on the characterization and engineering application of 3D printed CF/PA composite was evidenced.

© 2020 Elsevier B.V.This study introduces fast manufactured (4−5 hours) vertical aligned carbon nanotubes/polydimethylsiloxane (VACNT/PDMS) stretchable sensor for aircraft structures and polymer composites. The VACNT growth method from thermal CVD with camphor and ferrocene precursors create a dense and homogeneous CNT structure with many ohmic conducting paths that surpass any more resistive tunneling phenomena. The extensive AC/DC piezoresistive performance investigation shows ohmic conduction. The VACNTs/PDMS sensor presented high linearity (r2 = 0.99) under ∼20 % strain with an average gauge factor of 3.28 at DC measurements and minor resistance variations (0.052 %) attributed to the copper terminals electrodeposition. An important feature is the resistance value compatible with conventional extensometry equipment. The DC gauge factor was ∼60 % higher than the conventional metallic strain gauge allowing measurements within a wider strain range.

© 2020Aeronautic structures are exposed to a great variety of temperatures and humid environments during service. Nowadays, it is known that the combined influence of moisture and temperature induces further detrimental effects on the fatigue behavior of bonded joints when compared to the influence of each isolated condition. The effect of hygrothermal pre-conditioning on the fatigue delamination growth onset of different bonded technologies was investigated. This paper provides a material database for composite-joints tested under different environments and gives important insights on the failure mechanisms observed under cyclic loadings. This information can be latter used to validate analytical and numerical models.

© 2020 Elsevier LtdRecent aircraft are increasingly presenting unconventional wing configurations, resulting in unusual aeroelastic responses and consequently giving rise to different technological strategies to enhance aeroelastic stability. Among them, the technique of stress stiffening by piezoelectric actuation emerged as a promising technological solution to improve the aeroelastic stability of structures with both ends axially constrained. Therefore, an aeroelastic model employing smart composite beam elements and time domain aerodynamic loads with strip theory for stress stiffening aeroelastic problems was developed and carefully validated. In addition, the effect of bending-torsion coupling provided by concentrated masses on the aeroelastic response of the structure is also taken into account, which was included by the presence of a slender ballast arbitrarily positioned along the span and chord. Parametric studies were performed investigating the influence of aspect ratio, fiber orientation angle, ballast position, as well as piezoelectric unit position along the span and its input voltage. Results showed a promising performance of such technique, as it could increase the bandwidth of two flexible modes associated with the flutter mechanism.

© 2020 Elsevier LtdA cohesive zone model is proposed to analyse composite delamination propagation under high-cycle fatigue loading. A new method to compute the strain energy release rate at any point within the element fatigue life cycle range is presented. The proposed scheme is based on the J-integral method evaluated at an estimated position of the crack-tip within the element, instead of the element integration points. Furthermore, a new fatigue damage evolution law is proposed to account for the unwanted quasi-static damage during the element fatigue degradation process. The model prediction capabilities were verified against experimental data available in the literature and theoretical solutions using a double cantilever beam configuration for mode I loading, four-point end-notched flexure configuration for mode II loading, and mixed-mode bending configuration for mixed-mode loading. The simulations were performed at both constant and variable amplitude loading. The numerical predictions obtained using the proposed model correlated very well with literature's experimental data.

© 2020 Elsevier LtdAdhesively bonded composite structures, if designed properly, have proven to be stiffer and to possess a higher specific strength than their mechanically fastened counterparts. To increase the applicability of these bonded joints in the aircraft industry, a study was performed to investigate the influence of hygrothermal aging on co-bonded composite stiffened panels with an initial disbond under cyclic compression loading. Experiments showed that hygrothermal aging led to a decrease in disbond growth throughout cyclic loading. The decreased disbond growth was likely caused by the increased ductility of the bond due to the presence of moisture. A higher ductility can lead to crack blunting and stress relaxation, resulting in higher fracture toughness of the bond. Furthermore, it was shown that hygrothermal aging did not influence the residual strength and stiffness of the panels after cyclic loading. The experiments were simulated numerically to gain a better understanding of the crack growth behavior and to aid future numerical crack growth predictions.

© 2020 Elsevier LtdMode I interlaminar fracture toughness of CFRP joints co-cured and co-bonded was experimentally characterised under several environmental conditions. Two test campaigns were carried out: one with as-received specimens tested at −54 °C and 25 °C, and another with hygrothermally aged samples tested at 25 °C and 80 °C. Dynamic mechanical analysis and scanning electron microscopy were used to explain distinct results. For co-cure, the propagation onset for as-received samples started at the interlayer region between the two adherends before migrating deeper into the adherend, while for as-received co-bonded samples, propagation onset was mainly cohesive. After aging, both bonding techniques failed directly in the adherend, in a fibre-tear fashion. Thus, the aging process presented a higher influence on failure mechanisms than the testing temperature. Regarding performance, mode I fracture toughness did not present a great variation for co-cured samples, since failure locus was always in the adherend. However, the initiation mode I fracture toughness of co-bonding decreased between as-received and aged samples. This difference is attributed to the presence of the adhesive in co-bonded systems, which showed to be more affected by environmental conditions than systems formed by fibre and matrix only, such as co-cured ones.

© 2020, The Brazilian Society of Mechanical Sciences and Engineering.A constitutive damage model is proposed in order to investigate the crushing response of 0 ∘ plies composite laminates, using a VUMAT subroutine implemented in ABAQUS/Explicit. This damage model predicts five failure mechanisms commonly observed in unidirectional carbon fibre-reinforced composite structures: fibre failure in tension, fibre failure in compression, matrix cracking in tension, matrix cracking in compression and in-plane shear failure. Its formulation is based on an energy framework which combines stress-based fracture mechanics and damage mechanics approaches within a unified way, enabling the prediction of the five failure mechanisms aforementioned in terms of damage initiation and damage propagation. In this work, a new strategy is also implemented to remove fully damage elements in order to ensure numerical stability and avoid element distortion problems. An experimental test campaign is proposed for verification of the numerical models. Preliminary results have shown a fairly good correlation between numerical predictions and experimental results for wedge-shaped tip composite laminates.

© 2020 Elsevier LtdA discussion on cohesive zone model formulation for prediction of interlaminar damage in composite laminates is presented in this paper. The degradation of interlaminar mechanical properties is analysed from a physical point of view. Firstly, the damage evolution is evaluated according to the traction-separation law and it is demonstrated that if a linear elastic unloading/reloading curve is assumed, the softening function must also be linear. Secondly, issues regarding damage onset and fracture criteria in mixed-mode loading are critically addressed and commented. A new set of criteria is proposed, and the limitations of existing criteria are discussed.

© 2019 Elsevier LtdAdhesive bonding technologies exhibit several advantages over conventional mechanical fasteners which include lower weight, reduced stress concentration in the adherents and excellent fatigue properties allowing the design of smooth surface contour and damage tolerant aerostructures. The overall structural performance of the composite joint depends on several factors related to the manufacturing process, such as surface preparation procedure, adhesive type, aging effects, and deficiency of inspection procedures. For this reason, there is a clear need to understand better how the joints behavior can be affected by them and the causes of their failure in order to improve the design and performance. In this context, this work presents the application of Scanning Electron Microscopy (SEM) technique to perform the fractographic analysis in two different types of joints namely co-bonded (CB) and secondary bonding (SB) that were submitted to a interlaminar fracture toughness tests under Mode II loading at Room Temperature Ambient (RTA) and Environmental Temperature Wet (ETW) conditions. The fracture surface of each sample submitted Mode II was prepared and analyzed regarding its failure aspects, enabling their association with the mechanical behavior of the samples during the fracture toughness tests and the values obtained. A detailed study on the fracture aspects of the composite joints was carried out correlating the fracture aspects with the mechanical behavior, type of processing and environmental conditioning in which each type of joint was tested, thus validating the use of these joints for operating conditions similar to those experienced in-service concerning aeronautical applications.

© 2019 Elsevier LtdAdhesive joints exhibit several advantages over conventional joints based on mechanical fasteners such as more uniform stress distribution, enhanced fatigue performance, stiffer connection, lower weight, smooth surface countor. However, the influence of environmental effects related to temperature and moisture absorption on the mechanical performance of these types of composite joints are not yet fully understood. This work investigates the hygrothermal effects on fracture toughness of composite carbon/epoxy joints under Mixed Modes I/II loading in different mode ratios (35%, 50% and 75%). Joints were produced using co-curing (CC), co-bonding (CB) and secondary bonding (SB) technologies. The specimens were submitted to an elevated temperature wet condition (ETW) at a high moisture content (90% R.H.) and high temperature (80 °C) and compared with results obtained at a room temperature ambient (RTA). The SB samples under ETW condition exhibited higher GII/GT values when compared to CC and CB specimens tested under the same aging condition. Using the scanning electronic microscopy (SEM) technique, it was possible associate the delamination process behavior of aging specimens with the fracture toughness values. The wet and hot environment affect both adhesive and adherent what results in significant changes in the failure aspects during the delamination induced failure process.

© Springer Nature Switzerland AG 2020.This paper presents a novel damage mechanics based failure model enabling the prediction of low cycle fatigue life and residual strength of isotropic structures under multiaxial loading. The approach herein proposed does not discretize every load cycle but instead takes an envelope loading whereby the numerical load remains constant at a maximum load level and the number of cycles is obtained from a given elapsed time defined within a pseudo-time framework. The proposed formulation is based on the smeared cracking approach accounting for damage propagation due to static and fatigue loadings, where the static component is based on the Von-Mises yield criterion and Prandtl-Reuss stress flow rule; whereas the crack propagation in cyclic loading component is based on the Paris-law. Furthermore, the formulation combines damage mechanics and fracture mechanics within a unified approach enabling the control of the energy dissipated in each loading cycle.

© 2019 Society of Plastics EngineersAn experimental investigation of mode-I translaminar fracture toughness and fatigue crack growth behavior of a carbon fiber-epoxy plain weave laminate manufactured by resin infusion under flexible tooling (RIFT) is presented in this article. Pre-cracked compact tension (CT) specimens were used to perform both quasi-static and fatigue tests. Different data reduction techniques for fracture toughness calculation were used and compared with each other. The ASTM E399 test method was modified to account for the material orthotropy and specimen geometry effects using a correction function based on a numerical evaluation of the strain energy release rate. The proposed modification shows good agreement against other experimental methods found in the literature and its application was validated for fatigue tests. Fatigue testing shows that failure in undesired modes is likely to occur prior to translaminar fracture, which was attributed to a higher tensile fatigue threshold than compression or shear fatigue threshold presented by the composite in analysis. Increasing the specimens’ initial notch length was a solution for avoiding these types of failure. The experimental results were compiled in the form of a Paris curve, and their particularities were discussed. A fractographic analysis was carried out to define damage patterns and its evolution process in both types of tests. POLYM. COMPOS., 40:3791–3804, 2019. © 2019 Society of Plastics Engineers.

© 2018 Society of Plastics EngineersAdhesive joints are being more extensively applied in the aeronautical industry, allowing for better integration between the structural parts and overall lower weight if compared with joints made with fasteners and rivets. However, a further evaluation of these new technologies is needed, once their critical fracture toughness under different environmental conditions is still unknown and this value is essential for design and certification of aircraft manufactured with these materials. Thus, this work focuses on the mode II fracture toughness characterization of carbon fiber composite laminates joined by co-bonded (CB) and secondarily bonded (SB) techniques, using EA 9695 epoxy adhesive aged at different environmental conditions (room temperature ambient—RTA and elevated temperature wet—ETW). Dynamic mechanical analysis (DMA) was used to understand the effect of moisture absorption on the glass transition of materials and on the decreasing of Mode II fracture toughness after aging. The DMA results showed a reduction of 11% in Tg values for the GIIc values of ETW samples in comparison with specimens tested at RTA condition. Reductions about 92 and 94% in Mode II fracture toughnesses were obtained for CB and SB aged specimens, respectively when compared with the toughness values obtained for specimens tested at RTA. Further inspection of the fracture surfaces using scanning electron microscope proved that light fiber-tear fracture occurred at both RTA and ETW conditions for CB joints, while fracture was mainly light-fiber-tear at RTA condition, becoming mostly cohesive after aging for SB joints. POLYM. COMPOS., 40:3220–3232, 2019. © 2018 Society of Plastics Engineers.

© 2018 Society of Plastics EngineersAdhesive bonding technologies for thermosetting polymer composites have been applied in several industrial sectors, such as marine, automobile, construction, and aeronautical industries due to their excellent mechanical behavior over conventional joining methods. One of the main drawbacks of this joining technology is that they are prone to delamination whilst in service. The overall composite joint structural performance depends upon several factors related to the joint manufacturing process such as surfaces preparation procedure, loading condition, adhesive type, aging effects, and inspection procedures. For this reason, there is clear need to better understand how the joint behavior is affected by these factors and the failure causes in order to improve the design and the joint performance. Within this context, this work presents a comparative fractographic analysis for two different joints types named co-bonded and secondary bonded tested under Mode I delamination at room temperature. One sample of each was observed through fractographic analysis, to identify similarities and differences between the fracture aspects, which may explain differences in toughness values and fracture behaviors. The main contribution of this article is a new failure analysis methodology focused on a better understanding on failure characteristics of adhesive joints providing a deep and critical insight into main failure mechanisms and damage sequence in composite bonded joints, which may aid analytical validation and numerical models for this type of joints developed elsewhere. Although fracture toughness (GIc) values of each joint are quite similar, some failure aspects and adhesion mechanisms differentiate these two bonding technologies. POLYM. COMPOS., 40:2973–2983, 2019. © 2018 Society of Plastics Engineers.

© 2019 Elsevier Masson SASAmong the aeroelastic phenomena most commonly affecting flexible and very flexible aircraft, those caused by gusts deserve special attention due to their potential either in degrading flying qualities and ride comfort or in increasing structural loads. It is then of interest to structural loads and flight controls engineers that solutions be developed to attenuate the effects of gusts on aircraft. Particularly, the use of piezoelectric transducers arises as one of the potential solutions in the design of gust load alleviation and structural mode suppression systems. In this paper, the gust load alleviation on a flexible smart idealized wing using only piezoelectric transducers is analyzed and experimentally tested. The numerical model includes a finite-element model of the wing, employing two-node, seven-degree-of-freedom smart beam elements, assuming small deformations and neglecting transverse shear. A quasi-steady, strip-theory-based aerodynamic model is used. Two control laws are evaluated: one based on output feedback, and the other based on feedback of observed states of a truncated system. Using a gust generator, wind-tunnel tests were performed at different flow speeds and gust frequencies to validate the computational model and to verify the performance of piezoelectric transducers. The results show a considerable attenuation of the wing root bending moment, especially using two piezoelectric actuators. Important performance improvements were overall verified with feedback of observed states when compared with static output feedback, specially to decrease the participation of the elastic modes in the gust response.

© 2018 Wiley Publishing Ltd.Composite joints exhibit different behavior in regard to delamination resistance when dealing with fatigue phenomenon. This research work focuses on an investigation to understand the failure mechanisms on the interfacial strength domain for delamination onset in cocured and cobonded joints. The analysis was based on strain energy release rate versus number of cycles plots that were obtained from fatigue tests in mode I with a stress ratio R = 0.1. The analysis encompassed from the microscopic to mesoscopic level obtained from scanning electron microscopic, and the images processed to extract the most relevant fracture patterns. The main difference between the two technologies was the stress concentration at the crack tip in which the cobonded joint presents a fabric carrier that blunts the adhesive layer, then delaying the delamination. This paper provides important information and guidelines to aid designers in the selection of the best composite joint for high-performance structural applications.

© 2018 Society of Plastics Engineers This work investigates and compares the interlaminar fracture behavior of composites manufactured by vacuum-assisted resin transfer molding subjected to three different temperatures (–54, 25, and 80°C) and mode ratios (25, 50, and 75%). The results indicate ductility enhancement with increasing temperature, which were confirmed by fractographic analyses. In tested specimens with 25% mode ratio, the G I and G II values were not greatly affected by the temperature. As the temperature and the mode ratio increases, the specimens exhibited higher G I and G II values compared with those measured at −54°C. In the tested specimens with 75% mode ratio, an unstable crack propagation was observed at −54°C due to brittle behavior of matrix, which is promoted by the decrease of the adhesion represented by cusps and broken fibers in SEM images. The cusps formation is less pronounced for specimens tested at 80°C and the fracture surface is flatter compared with those tested at −54 and 25°C. POLYM. COMPOS., 40:E1029–E1040, 2019. © 2018 Society of Plastics Engineers.

© 2018 Elsevier LtdA finite element model for predicting the nonlinear aeroelastic behavior of composite panels undergoing intralaminar and translaminar progressive damage in supersonic flow is presented. The classical plate theory in conjunction with the von Kármán nonlinear strains is used for structural modeling, and the linear piston theory is used to model the aerodynamic loads. Progressive damage is modeled by a smeared cracking formulation in which stress-based, continuum damage mechanics and fracture mechanics approaches are combined. No modal reduction is performed and an iterative form of the Newmark method is used for the numerical direct integration in time of the nonlinear equations. Simulations considering different lay-ups are conducted, in which the influence of progressive damage on the aeroelastic behavior of the panels is investigated, and damage extent and failure mechanisms are assessed. The results obtained in the analyses consist in important insights concerning the flutter-induced damage in composite panels.

© 2018 Elsevier LtdThe increasing need for automatic mesh generation has led to the development of efficient triangulation algorithms that are able to discretize any 2D or 3D domain. Modern finite element formulations based on strain smoothing techniques (SFEM) provide enhanced convergence properties, preventing yet the stiffening behavior of triangular meshes. Recent research has shown that meshless methods based on triangular mapping of the integration domain can be used to produce even better convergence properties than SFEM. The present study explores the Edge-based Smoothed Point Interpolation Method (ES-PIM) as a meshless solution to investigate linear buckling on variable angle tow (VAT) laminates. Such advanced composite structures show a heterogeneous distribution of constitutive properties and thickness, presenting additional challenges to the numerical solution. Important aspects related to the transverse shear correction herein adopted are investigated, leading to interesting conclusions regarding the possibility to use the ES-PIM for conservative estimates of the critical buckling load of VAT laminates.

© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.The effect of addition of viscoelastic layers on the acoustic scattering quadrupoles near the trailing edge of composite plates is evaluated. For modelling of the viscoelastic material the complex modulus approach was used in combination with the frequency-temperature correspondence principle. The computation of laminate stiffness is based on Classical Lamination Theory. We employ a numerical method to compute the acoustic field scattered by finite elastic plates. Based on a Boundary Element Method, this procedure solves the Helmholtz equation subject to boundary conditions related to the vibration of the plate. These conditions are recast in terms of the vibration modes of a rectangular plate. Results show that by adding viscoelastic plies to a composite plate we modify the far-field sound scattered by turbulence near an edge of the plate. Parametric studies show that this approach reduces scattered noise at resonance frequencies. Discussions on the operating temperature, positioning and thickness of the viscoelastic layers are provided.

© Universal Technology Corporation, 2018.This work presents the development of a finite element beam model accounting for structural and aerodynamic nonlinearities regarding large deflections. The total Lagrangian formulation is employed for describing the exact Timoshenko’s bending kinematics, whereas the torsion is modeled as an uniform torsion and uncoupled from the bending motion. The aerodynamic description is based on an unsteady 2D strip theory in the time domain with the Jones approximation for Wagner’s function. In addition, a follower forces assumption combined with a simplified stall model are assumed, in which the lift-curve slope is interpolated based on the experimental data available in the literature. Very good correlation between experimental and predicted aeroelastic responses has been obtained for a highly elongated plate-like wing structure with a ballast at the free end.

© 2019 The Author(s) Published by the Royal Society. All rights reserved.We investigate the effects of structural damping on the interaction of a turbulent eddy with flexible plates with respect to the efficiency of aerodynamic noise generation. Potential benefits are studied using a model based on a point-reacting compliant semiinfinite plate on a spring-damper foundation. This scattering problem is solved using the Wiener- Hopf technique. We compare results for semi-infinite compliant plates with finite ones. In both cases, plate vibration lead to reductions of sound radiation, especially at resonance; damping tends to reduce such acoustic benefits. We also present a formulation that considers the effect of structural damping on the acoustic properties of finite elastic plates. Numerical results are obtained by applying a boundary element method to solve the Helmholtz equation subject to the boundary conditions imposed by the plate vibration. Under specific conditions, such as high fluid loading factor and low bending-wave Mach number, the acoustic power scattered by an edge tends to be smaller than that which propagates over the plate as bending waves. Results show that structural damping attenuates these waves and may modify the far-field acoustic pressure, mostly by reducing the scattered sound at structural resonances. All models show that large damping coefficients lead to locally overdamped responses. There is thus an ideal range of structural damping to reduce both plate vibration and acoustic scattering.

© 2018 Elsevier LtdThis work focuses on Mode I fatigue induced delamination growth onset characterization of co-cured (CC), co-bonded (CB) and secondary bonded (SB) composite joints with an epoxy interleaf. The CC joints used in this work comprise only fibers and resin, while, CB and SB joints contain an adhesive film. Experimental tests were carried out at room temperature using double cantilever beam (DCB) specimens. The delamination behavior was evaluated in terms of the strain energy release rate (SERR) considering a no-growth criterion based on 106 cycles without crack propagation. For Mode I cyclic loading, the threshold values of the CB and SB joints were about 2.5 times higher than the one achieved by the CC joints, presenting a better performance in terms of fatigue delamination growth-onset. A fractographic study was conducted using scanning electron microscopy in order to relate the SERR results with the joint′s failure mechanisms. It was observed that there is a good mechanical compatibility between the substrate and adhesive in the CB and SB joints studied herein. The results found in this paper indicated that the Mode I SERR threshold values were not significantly affected by the interleaf.

© 2018 Society of Plastics EngineersAn experimental study on mechanical properties, such as strength, elastic modulus and interlaminar fracture toughness of two different manufacturing processes namely autoclave vacuum bagging (AP) and resin liquid Pressurized Prepreg (PP) is outlined in this article. AP is a technique employed to create mechanical pressure on a laminate during its cure cycle. This process have for years enabled aerospace industry to maximize the physical properties of advanced composite materials. The resin liquid PP is an out-of-autoclave process in which prepregs are laid up onto a closed-mold, liquid resin is injected to pressurize the laminate in order to exert the hydrostatic pressure required to consolidate the preform and eliminate any gas bubbles that may form during the setting of the resin and the cure is performed into a heated press. The experimental results indicated that laminates manufactured by the PP process exhibited an overall inferior mechanical performance when compared with the laminates obtained using the AP process. The reduction in the mechanical properties may be explained by the non-homogeneous resin perculation, interfacial mixing of two types of resins, higher thickness and, consequently, higher resin content of the PP laminates. POLYM. COMPOS., 39:E2562–E2572, 2018. © 2018 Society of Plastics Engineers.

© 2018, ASM International.The aeronautic structures normally operate under high levels of hygroscopic moisture from the surrounding environment at different temperature ranges while in service. Under such conditions, the behavior of laminate composite submitted to cyclic or static loadings can change drastically. In order to understand those effects in stitched fabrics, fatigue tests with open-hole specimens were carried out with a stress ratio of R = −1 and R = 0.1. The specimens were fatigue-tested as provided (environmental conditions) and after exposed to hygrothermal weathering conditions. Based on evidences from recent studies available in the open literature, e.g., effect of water diffusion on epoxy matrix, the overall results indicated a significant reduction in stiffness after the specimens are exposed to hygrothermal effects. The reduction in matrix stiffness, in this case, enhanced the fatigue strength in tension–tension load (R = 0.1) when compared to the specimens in normal conditions. The opposite occurs for the specimens loaded with stress ratio of R = −1, in which the delamination mechanisms changed during the loading reversion from tension to compression that promoted early delamination. Therefore, this process reduced the fatigue life of the specimens under hygrothermal condition. Then, by fractographic investigation, it was verified fracture patterns that regard to mode II damage in R = −1, in which mode II fracture toughness is known to decrease in the presence of water molecules.

© 2018 Elsevier LtdComposite materials undergo intricate damage processes, which are accentuated when these materials are exposed to impact loading conditions. In this context, this work aims to study the behavior of fiber reinforced polymer composites submitted to high strain rate in compression. A composite laminate plate was obtained using a plain weave carbon fiber fabric and an epoxy resin as matrix. The Split Hopkinson Pressure Bar (SHPB) technique was used to apply compressive loads at three different strain rates and in six different directions (0° 15° 30° 45° 60° and 75°) relative to the warp. The compressive failure modes were studied using a high speed camera to record the SHPB test. With increasing strain rates, it was possible to identify a transition from longitudinal cracking failure to delamination buckling failures. Similarly, the off-axis loading conditions (mainly for 45°) resulted in extension-shear coupling effects, which promoted the delamination buckling failures.

© The Author(s) 2018.Vacuum assisted resin transfer molding is a promising process in advanced composite manufacturing with a wide range of applications in industry. That potential is often misused, though, because of the lack of an efficient and reliable simulation tool to support product development. Most of the simulation methods in use today are based on Darcy’s law, which explains the permeation of a fluid in a porous medium. However, it is known that this law has limitations when applied to the context of dual-scale fibrous reinforcements: macro porosity given by fiber architecture generates resistance to flow, while the inner porosity inherent to fiber tows causes it to absorb resin, affecting the flow. The latter effect cannot be explained by traditional theory. In order to explore these limitations, this work proposes a simplified model to vacuum assisted resin transfer molding process from the point of view of system dynamics, and to prove the viability of such theory. The ultimate goal is to propose a more complete model in light of system dynamics that saves time and cost while offering the same reliability as current simulation models. In order to provide an explanation to both dual-scale phenomena, a parallel association between a resistance and a fluid capacitance is proposed. Model validation is then performed through the analysis of experimental data followed by the comparison between the Darcy infusion profile and the one predicted by the resistor-capacitor-parallel (RC-parallel) circuit model. Thus, this work is able to perform a proof of concept that leads to a novel and yet unexplored field of study.

© 2018, Brazilian Association of Computational Mechanics. All rights reserved.Shape distortions and warpage are a major source of problems for composite manufacturers. These distortions are usually accompanied by built up residual stresses. They can deform a component so that it becomes useless. It also has the capability to reduce the strength of the structure. In this paper, the three-dimensional version of the constitutive model originally proposed by Svanberg and Holmberg is employed to predict the warpage of a wing planform. The model takes into account important mechanisms such as thermal expansion, resin shrinkage and frozen-in strains developed during curing cycles. The model was implemented into ABAQUS Finite Element code as a user subroutine UMAT. The macromechanical properties of each composite layer were predicted using a micromechanics based approach, implemented into MATLAB. Results show that wings with cross ply laminates with reducing thickness along the span experienced more warpage than quasi-isotropic laminates. Furthermore, for wings with equal thickness along the span, the results show that the quasi-isotropic laminates experienced more warpage than cross ply laminates. Lastly, the results show that wings with progressively reducing thickness experience twist that is varying from the wing root to the wing tip while wings with a constant thickness experience twist mainly at the centre of the wing.

© 2018 Elsevier LtdThis paper presents a numerical and experimental study on the intralaminar tensile fracture toughness of carbon fiber reinforced composite subjected to high strain rates. As there is no standardized testing procedures for intralaminar fracture toughness characterization of composites at high strain rates, there is a clear need to design specimen geometries, testing apparatus and data reduction schemes that allows the characterization of the fracture toughness of composites in the dynamic regime. Initially numerical studies were performed based on finite element simulations in order to investigate the viability of its construction for different testing configurations to characterize the intralaminar toughness of composite laminates. A comparative study is presented showing the advantages and disadvantages of each testing configuration. A new data reduction scheme based on modifications in the ASTM standard, accounting for material anisotropy and specimen finite geometry effects is suggested. Experimental tests were carried out, using the proposed specimen configuration at different strain rates in order to investigate the strain rate effects using a modified version of the Split Hopkinson Pressure Bar. Fractography analyses using Scanning Electron Microscopy(SEM) have been also performed in order to investigate the strain rate effects on the failures mechanisms of the composite material studied herein.

© 2018 Acoustical Society of America.Trailing edge scattering is a significant source of sound, and elasticity is known to decrease the radiated sound by a process involving coupled acoustic and bending waves. Most of the analysis available in the literature to deal with this problem is limited to structures of isotropic material. A numerical method is extended, based on the solution of a boundary element method with boundary conditions given by the structural problem, to account for anisotropic composite plates, restricted to symmetric laminates. These conditions are recast in terms of the vibration modes of a rectangular plate. To obtain these modes, the hierarchical finite element method is used to model an elastic flat plate. Expressions for bending waves propagating in such plates are derived, and how the solution of the problem is modified to account for these effects is shown. Results show modifications in the scattered sound as a function of ply orientation and stacking sequence. Composite materials are shown to be advantageous, since laminates lead to lower acoustic scattering when compared to structurally equivalent metallic plates. This is due to a lower specific mass, leading to higher coupling between fluid and solid, and thus to more significant elasticity effects, decreasing substantially the radiated sound.

© Published under licence by IOP Publishing Ltd.This paper analyses the Mode II fatigue delamination growth onset for secondary bonded joints compared to co-cured joints. The materials used were composed by two carbon fiber reinforced sub-laminates joined by the secondary bonded and co-cured (without adhesive) methods. Mode II fatigue tests were performed using three-point bending End Notched Flexure test setup. The tests were performed under displacement control and a sinusoidal displacement applied at a frequency of 5 Hz with a Rd = 0.1, that mean, δmin = 0.1 δmax. The main objective of this study was to obtain the strain energy release rate (SEER) versus number of cycles (Nf) in order to evaluate the effect of the adhesive on fatigue life of bonded joints. A Closed form solution for 3-ENF setup was used in order to define the displacement amplitudes that were applied in the fatigue test. The results show that the use of adhesive causes a reduction on Mode II fatigue delamination growth onset SERR (Gth) for secondary-bonded compared to co-cured joints. Finally, a scanning electron microscopy (SEM) was used to analyze the fracture surfaces of the Mode II secondary bonded specimens. The fractography images show no crack growth for the specimens tested below or at Gth loading levels, which indicated that the value corresponds to the threshold.

© 2017, The Author(s) 2017.The main concern related to the flutter phenomenon is predicting and avoiding it. This paper describes the application of a flexural-torsional flutter testbed for acceleration reduction by applying active and passive model-based control. The model consists of the 2D typical section, with aerodynamic loads estimated by an unsteady time-domain formulation based on Wagner’s function. The active control architecture consists of a stability augmentation system with output feedback and gain scheduling via the linear-quadratic regulator theory and actuation by servomechanism. The passive control employs a shape-memory alloy to provide additional torsional stiffness. Experimental results show considerable reduction of oscillations at a relative low cost for both active and passive control strategies, and that the use of shape memory alloys in aeroelastic stability problems is promising.

© 2018, The Brazilian Society of Mechanical Sciences and Engineering.The objective of this work is to use an asymptotic homogenization numerical model to obtain elastic properties of unidirectional fiber-reinforced composites. Square and perfect hexagonal unit cells are employed, and the influence of the fiber volume fraction over the homogenized elastic properties is studied. The effectiveness of the predictions is assessed by comparisons to experimental properties, and also another micromechanical model based on a representative volume element. The composites E-Glass 21xK43 Gevetex (glass fiber)/LY556/HT907/DY063 (epoxy matrix) and AS4 (carbon fiber)/3501-6 (epoxy matrix) were studied and good agreement between experimental and numerical predictions was found. Discrepancies between experimental and numerical data are explained in terms of simplifications considered in the homogenization model. An adjustment of properties here performed, based on varying fiber volume fractions, showed to be effective to physically represent the studied fiber composites in a micromechanical stress model based on asymptotic homogenization, developed to estimate failure envelopes of such materials.

© 2017, © The Author(s) 2017.This paper shows a developed process to reclaim carbon fiber from end-of-life thermoset composite or pre-preg process waste, which uses pyrolysis and oxidation to remove the matrix (resin) and a plasma reactor to treat the exhaust gases. Laminates were manufactured to be recycled and the reclaimed laminates were remanufactured and then tested. Tensile tests, interlaminar shear strength tests and measurement of the fiber volumetric fraction for both virgin and reclaimed laminates were carried out. The dimensions, masses and permeability were also measured for both virgin and recycled laminates. Additionally, ultrasound inspections, Raman spectroscopy, micrographs using scanning and transmission, as well as microscopy of the fracture surfaces of the composite specimens submitted to the tensile and interlaminar shear strength tests were performed. Monofilament tensile tests in one of the reclaimed and virgin samples were also conducted. All these tests and analysis were conducted aiming at comparing the overall performance of the reference (virgin) composite to the one manufactured with reclaimed carbon fiber, trying to better understand the differences between them, and the origin and cause of these differences.

© 2017 Elsevier LtdThe flutter boundary prediction of complex aeroelastic systems is not an easy task. In some cases, these analyses may become prohibitive due to the high computational cost and time associated with the large number of degrees of freedom of the aeroelastic models, particularly when the aeroelastic model incorporates a control strategy with the aim of suppressing the flutter phenomenon, such as the use of viscoelastic treatments. In this situation, the use of a model reduction method is essential. However, the construction of a modal reduction basis for aeroviscoelastic systems is still a challenge, owing to the inherent frequency- and temperature-dependent behavior of the viscoelastic materials. Thus, the main contribution intended for the present study is to propose an efficient and accurate iterative enriched Ritz basis to deal with aeroviscoelastic systems. The main features and capabilities of the proposed model reduction method are illustrated in the prediction of flutter boundary for a thin three-layer sandwich flat panel and a typical aeronautical stiffened panel, both under supersonic flow.

© 2017 Elsevier LtdThis paper presents an experimental and numerical study focused on the mode-I intralaminar toughness characterization of a woven carbon/epoxy composite loaded in compression and subjected to high strain rates. Simulations for non-standardized Single Edge Notch Bending (SENB) and Double Edge Notch (DEN) specimens were carried out using a continuum damage mechanics based failure model implemented as an user defined material model within ABAQUS software. A Finite Element Model was used in order to produce an optimal specimen for intralaminar fracture toughness tests. A new data reduction scheme based on the numerical evaluation of the strain energy release rate using the J-integral method is proposed to determine the stress intensity factor for composites. The proposed methodology accounts for finite geometry and material anisotropy effects. The dynamic tests were carried out at strain rates of 560s-1,690s-1,770s-1 using an adapted version of the Split Hopkinson Pressure Bar. A high-speed camera was used for monitoring the crack propagation. A Scanning Electron Microscope (SEM) was used to aid the fractographic analyses on the damaged surface of the tested samples searching for the possible failures mechanisms within the material. The experimental results indicated that the composite laminates studied herein are very sensitive to the strain rate effects.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.The use of morphing structures aims to increase aerodynamic efficiency, decreasing fuel consumption and aircraft overall weight. Within this context, piezoelectric materials are of great interest in the design of smart structures because the piezoelectric effect is reversible, allowing them to work both as sensors and as actuators. The present work proposes a geometrically nonlinear finite element formulation for composite beams with embedded piezoelectric layers for application in morphing aerostructures. The formulation uses the complete Green strain tensor to account for geometric nonlinearities, and linear piezoelectricity to model the electromechanical behavior. A set of nonlinear equilibrium equations results from the application of variational principles, and is finally solved by means of an iterative-incremental arc length method. The sensitivity of the element to stacking sequence and number of actuators are investigated.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.Three numerical models for the prediction of high-cycle fatigue-driven delamination in carbon/epoxy composite laminates are compered through finite element analysis of a mixed-mode bending specimen. These models allows delamination modelling without knowing, a priori, the modes ratios. The models are implemented into ABAQUS/Explicit FE code within solid elements. The local part (at element level) of the algorithms are implemented in a user-defined material subroutine (VUMAT) and the non-local part (at structure level), in a VEXTERNALDB subroutine. The study found that the use of strain energy release rate based on Paris' law variation combined with cohesive zone model is a robust approach. The use of strength-based damage parameter to account for the fatigue damage reduces the numerical integration error associated with the cycle jump and the non-local crack tip tracking algorithm improves the accuracy.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.In this work we deal with the problem of trailing edge noise scattered by a flat elastic plate. We use a model based on a boundary element method that couples the acoustic problem with the fluid-structure interaction and takes into account structural damping. The solution is obtained using the modal basis of the free vibration problem. The objective of this paper is to expand the acoustic scattering analysis for different damped plates to increase knowledge about the effects of structural damping and to identify potential benefits of using inherently damped structures, such as viscoelastic materials. It is found that there is a range of damping coefficients, capable of reducing peaks in the acoustic spectra associated with structural resonance, while mantaining the reduction of scattered sound due to elasticity. When the damping coefficient is increased above this range, the rigid-plate limit is recovered and acoustic benefits are reduced. The present results allow the selection of optimally-damped structures with respect to acoustic radiation.

© 2018 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.The proximity of the source and an edge can make the acoustic scattering by wings a significant source of aerodynamic sound. Theoretical results have shown that elastic edges lead to reductions of acoustic scattering; however, experimental confirmation of theoretical trends is difficult, since surface vibrations modify both the source structure and the scattering properties. A simplified, controlled setting for measurements of acoustic scattering, allowing the evaluation of fluid-structure interactions, would thus be desirable to study how elastic edges modify the radiated sound. We present an experimental procedure to isolate the scattered field using a loudspeaker in the vicinity of at plates. The methodology is applied to three different plates, made of steel, aluminum and carbon fiber, as a demonstration. The responses of these elastic plates are studied for a sound source of dipole type near the trailing edge. The method is based on the experimental determination of frequency response functions between source and radiated sound for experiments with and without the plate; subtraction of results, accounting for amplitude and phase, isolates the scattered field. Experimental results treated with the developed procedure were compared with predictions made by numerical simulations performed with a Boundary Element Method (BEM), coupling the acoustic problem with the plate vibration. The comparison between experimental and numerical results revealed that a two-dimensional model can predict satisfactorily the reductions in scattered field by elastic plates observed in the experiment. The present methods can be used to support the choice between different materials for edges focusing on their respective acoustic benefit.

© 2017 Elsevier LtdThis work investigates the effects of temperature in the shape memory alloy hybrid composites (SMAHC) cylindrical stiffened panels’ aeroelastic stability. The SMAHC is modelled using a micromechanical formulation embedding carbon fiber, SMA wire and resin to the same lamina and taking into account the martensite/austenite phases of transformation in the material response. Virtual work principle formulation is implemented with classical laminate plate theory (CLPT) panel formulation and one-dimensional Euler-Bernoulli beam theory formulation for the stiffener. Numerical results are obtained by using an energy based semi-analytical method applying hierarchical polynomials to approximate the membrane and out of plane displacement fields. Different geometric configurations, laminate stacking sequences, boundary conditions and radii of curvature are investigated. The study shows that the variation of temperature induce stiffening due to changes in the martensite/austenite fractions of the SMA, increasing the critical flutter dynamic pressure. Therefore, it can be achieved certain control in the flutter critical boundary by increasing the temperature of the shape memory alloy (SMA) wire. The effects due to the SMA wire stiffening with the temperature are more pronounced for cross-ply stiffened cylindrical panels with unitary aspect ratio and for angle-ply panels with aspect ratio higher than one.

© 2017 World Scientific Publishing Company.The present paper describes a numerical modeling approach to predict impact resistance and residual Shear Strength After Impact (SSAI) of fiber reinforced polymer composites subjected to bird strike loading. An improved damage mechanics based on material model, previously developed by the authors, is combined with an equation of state to simulate the progressive failure in composite aerostructures subjected to bird strike loading. A series of bird strike impacts on flat panels fabricated from low cost woven glass composite materials are used to validate the material model for practical composite component applications. A numerical study on the residual SSAI of a typical composite shear web is also presented. The panels are modelled with shell elements only. The proposed material model formulation accounts for the strain rate enhancement to strength and shear nonlinearities observed in composite materials. A hydrodynamic model for the bird, based on 90% water and 10% air, is derived to represent the behavior of the bird for all impact scenarios considered. The bird is heterogeneous in nature. However, a uniform material behavior is assumed with a geometry based on a 2:1 length to diameter ratio with a cylindrical body and spherical end caps using Lagrangian mesh. Appropriate contact definitions are used between the bird and the composite panel. The simulations results are compared to experimental results and conclusions drawn.

© 2017 Wiley Publishing Ltd.Composite structures usually undergo to temperature variations in aircraft during landing/taking off and when cruising at high altitude. Under these conditions and in combination with curved structures, they can generate severe thermal stresses that induce delaminations. Considering the importance of studying delamination in these conditions, this research imposed an anti-symmetrical laminate to cyclic temperature variations of 130 °C and −70 °C with the objective of inducing varied curvatures and, consequently, crack growth. Different from standardized test procedures, this test setup elastically deformed coupons without external forces and forward experimentally and numerically evaluated the strain energy release rate (SERR) during crack propagation. This procedure enabled the assessment of delamination rate (da/dN) as a function of maximum SERR. The experimental results were compared with numerical results obtained by ABAQUS Finite Element code. Despite large scatter in experimental results, a reasonable correlation between experimental and numerical results was obtained in terms of crack growth rate (da/dN) as a function of the maximum SERR.

© 2016, © The Author(s) 2016.The increasing use of composite in the aircraft industry has raised the interest for a better understanding of the failure process in these materials, which can be also influenced by the manufacturing process of the laminate. Some materials used in vacuum assisted resin transfer molding process have been studied in the open literature but very few data have been published for resin transfer molding-6 epoxy based laminates, in particular studies showing the influence of the temperature on the interlaminar fracture behavior of this type of laminates. The aim of this article is to investigate the interlaminar fracture behavior of resin transfer molding-6 based carbon composite laminates manufactured by vacuum assisted resin transfer molding subjected to Modes I and II at 25℃ and 80℃. The results show the influence of the temperature on the interlaminar fracture toughness of composites and provide a database to design composite aerostructures subjected to temperatures commonly experienced in civil aviation. The fracture aspects of the tested laminates were also investigated and directly related to the trend in results found for the fracture toughness values.

© 2017, ASM International.In a 3D preform, the out-of-plane reinforcement is effective for decelerating or suppressing the delamination process as the non-crimp fabric does not connect the neighboring laminae effectively. Hence, the interlaminar strength of the stitched laminae is supposed to behave in the same way as a regular unidirectional composite. In order to determine whether or not the stitched yarns contribute to the interlaminar fracture toughness, this study determinated the delamination resistance of a quasi-isotropic laminate. The analysis was based on interlaminar fracture toughness (GIc) and propagation energy curve in tests conducted in mode I opening with double cantilever beam specimen geometry. The results of fracture toughness as well as strain energy for propagation were compared to their fracture surface. A decrease in the propagation energy prevailed in the surface because the stitch yarn replaced the carbon fiber/epoxy interface, which has better chemical affinities, i.e., covalent bonds.

© 2016 Elsevier LtdThe substitution of conventional mechanical fasteners by adhesive joints has been advocated by the aircraft and aerospace industries due to the weight saving potential. Flaws such as debonding of the adhesive layer between the skin and the stiffener may greatly affect the structural behavior of composite panels. Within this context, this work presents a semi-analytical approach for the numerical investigation on the effects of skin-stiffener bonding flaw size on the vibration and linear buckling behavior of T-stiffened composite panels. Skin and stiffener have been modeled using an assembly of curved and flat panel components, with each domain approximated using a set of hierarchical polynomial functions. A penalty-based approach has been used to assemble the various domains and to model the debonded region between the stiffener flange base and the plate. This approach ensures full compatibility in terms of displacements and rotations between the stiffener's base top face and the panel bottom face allowing to model different skin/stiffener debonding lengths. The results obtained using the proposed semi-analytical models have been compared and verified against numerical predictions based on finite element analyses.

© 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.A numerical method to compute the acoustic field scattered by finite perforated elastic plates is extended to include structural damping typical of viscoelastic materials. We employ a boundary element method to solve the Helmholtz equation subject to boundary conditions related to the vibration of the plate. In order to enable our investigation of the effect of damping, we rewrite the equations considering the terms responsible for the structural damping. Results show that by adding damping to the problem formulation, the flexural waves in the plate are attenuated and thus can modify the far-field sound scattered by turbulence near an edge of the plate. Parametric studies also show that structural damping tends to reduce scattered sound at structural ressonances. The combined effects of elasticity, porosity and damping may be more appropriate to represent the behavior of realistic materials.

© 2017 Universidade Federal de Sao Carlos. All rights reserved.Aiming to reduce aircraft weight, aeronautic industry seeks alternative materials and processes used to join its different structural parts. An option to traditional methods are high performance adhesive joints, which reduce weight, number of parts and component final cost, also resulting in higher strength structures. Although, the lack of experimental data to provide a detailed structural characterization of these joining techniques had limited their commercial application. The proposal of this work is to investigate the Mode I interlaminar fracture toughness under quasi-static loading using DCB specimens of carbon composite joints made by co-bonding and secondary bonding techniques, the latter giving more reliable results. For a better understanding on the failure in the systems, DSC and microscopy techniques were applied, from which three stages of delamination process during testing were observed: 1st Stage) Cohesive failure represented by an unstable crack propagation from a high energy level; 2nd Stage) transition from cohesive to adhesive and final intralaminar failure mode with lower energy levels than Stage 1; and 3rd Stage) completely stable propagation at low energy levels (delamination migrates from intralaminar to interlaminar, entirely in the substrate).

© 2016 Elsevier LtdThis work focuses on the determination of failure envelopes of unidirectional fiber reinforced composites. A two scale analysis is considered and the mathematical theory of asymptotic homogenization is applied to model the problem. For a given stress applied to the macro level, it is possible to assess stresses at the micro level domain. Three regions of the micro level are considered: matrix, fiber and the interface between them, and each region is ruled by its own failure criterion. A methodology to determine failure of composites using the homogenization is proposed. In the methodology, the strengths of the composite are used to determine the strengths of the constituents: a curve fitting adjustment is applied to calculate the strengths of the matrix and an analytical procedure is used to obtain the strengths of the fiber and interface. Then, the strengths of the constituents are used to evaluate failure criteria at the micro level, and the numerical failure envelopes are built. The advantage of the proposed methodology is that it is capable of calculating numerical failure envelopes with good approximation to experimental envelopes and also to the Puck & Schürmann criterion, requiring only five unidirectional strengths of the composite as inputs.

© 2016 Elsevier LtdA better understanding and improvements on the dynamic and aeroelastic behaviors of composite structures by using active and passive control strategies are nowadays key issues in designing advanced lightweight aerospace structures with smaller levels of vibrations in order to perform their tasks with success, reliability and safety. However, since light structures tend to be more flexible, it is necessary that the structure itself shows the ability of dissipating energy and stabilizing itself when subjected to external dynamic loadings imposed by the airflow. In this sense, smart materials can be used as an excellent alternative, being able to stabilize these structures. The interest here is to investigate the possibility of increasing the supersonic flutter boundary of a composite flat panel by applying a multimode shunted piezoceramic in series topology, in which active control strategies cannot be easily performed. Despite the fact that much research on passive aeroelastic control strategies have been conducted in the open literature, few works have been suggested the use of multimode shunt circuits to deal with the flutter problem of aeroelectromechanical systems, which motivate the study reported herein.

© 2016 Elsevier Masson SAS. All rights reserved.This work investigates the aeroelastic stability boundary of flutter in Shape Memory Alloy Hybrid Composite laminates (SMAHC). The SMAHC consists of SMAs wires and continuous carbon fibers embedded into a polymeric matrix resulting in a three constituent composite material. The derivation of the effective mechanical properties of the SMAHC is based on micromechanical model which accounts for temperature and fraction of martensite/austenite transformation phases of the shape memory alloy. Hamilton's principle is used for the formulation of the energy functional and to obtain the equilibrium equations and boundary conditions of the aeroelastic problem. The finite element method is employed to numerically solve the equations. Different geometric configuration, laminate stacking sequence, boundary conditions and curvatures are investigated. The study shows that the stiffening effect induced by the changes in the fraction of martensite/austenite transformation phases of the shape memory alloy increases the rate of occurrence of flutter, stabilizing the plate. Thus, one can control the occurrence of flutter speed by controlling the temperature of the SMA wires and the proper design of the geometric properties of the panel and tailoring of the composite laminate.

© 2016 Elsevier Masson SAS. All rights reserved.The present study involves the application of surface viscoelastic damping treatments to remedy panel flutter problems in existing aircraft components in which active control strategies cannot be easily performed. The rationale for such study is the fact that as the viscoelastic materials are often used to solve a variety of resonant noise and vibration problems in aerospace industry, it becomes important to quantify the increase of aeroelastic stability that can be obtained by the inclusion of viscoelastic treatments. The flutter boundaries of the aeroviscoelastic system accounting for the frequency- and temperature-dependent behavior of the viscoelastic material are computed by adopting the so-named Golla-Hughes-McTavish model. Since the inclusion of internal variables in the viscoelastic model leads to an augmented coupled system of equations of motion, a numerical pre-processing is found to be necessary prior to the resolution of the complex eigenvalue problem for the purposes of flutter analysis. After the theoretical foundations, the stability analysis of a three-layer sandwich plate under supersonic flow is addressed. The results show that it is possible to increase the critical flutter speeds of flat panels using surface viscoelastic damping treatments. However, the temperature and the thicknesses of the layers have significant effect on the flutter boundary.

© 2016 Elsevier Ltd.Flutter in aeronautical panels is a type of self-excited oscillation which can occur during supersonic flights. At the flutter point the vibrations of the panel become unstable and increase significantly in time. This manuscript presents a semi-analytical model taking into account the stiffener's base effects, in order to predict the aeroelastic response of laminated composite stiffened panels under supersonic flow. Krumhaar's modified supersonic piston theory, which considers the radius effect, is adopted to model the aerodynamic loading. The proposed model has been validated against results available in the literature for various configurations. A parametric study considering different panels and stiffener configurations is also presented. The numerical results indicate that the stiffener base significantly affects the panel aeroelastic behavior. Preliminary studies also indicate that redistributing the laminate plies from the stiffener's flange to its base significantly increases the torsion stiffness of the panel locally, opening new design possibilities that may lead to higher critical flutter speeds and therefore to better designs. The results also indicate that designs with plies distributed on the base may lead to a better flutter performance when the airflow is transverse to the longitudinal stiffener direction.

© SAGE Publications.This work presents an experimental characterization of the curvature effects on the compression-after-impact strength of laminated composite shells. Curved panels impacted on the outer (convex) face and with normal pressure on the inner (concave) face with three different curvatures at three different impact energy levels were tested. A compression-after-impact testing setup was designed and implemented to evaluate the impact-induced damage tolerance of the composite shells. An analytical modeling methodology for compression-after-impact strength predictions based on the Mar-Lin and Whitney-Nuismer failure criteria is also proposed. The approach proposed herein consists of replacing the damaged area of the impacted coupon by an equivalent hole. The analytical compression-after-impact predictions obtained using the Mar-Lin and Whitney-Nuismer failure criteria were compared with experimental results. A good agreement between analytical predictions and experimental results was found. The experimental results also indicate that the compressive residual strength of the composite shells is significantly affected by the shell curvature and internal pressure effects.

© 2015 Elsevier Ltd. All rights reserved.A major concern in laminated composites for structural applications has been crack nucleation between plies and its propagation through the interface. A better comprehension of stable crack propagation may lead to more reliable predictions of the rate with which cracks grow in weave fabric laminated composites. To this aim, a number of empirical relationships proposed in the literature were studied for their applicability to satin weave fabric (5HS) composite with spread tows. In order to understand the fatigue delamination process, Double Cantilever Beam specimens were submitted to cyclic loading, and their respective da/dN vs strain energy release rate (SERR) data collected and correlated to these empirical relationships. Despite that the empirical or semi-empirical relations fit equally well to the data, the well-known Hartman-Schijve equation was adopted. This equation was properly modified according to the fracture surface investigation which had its fracture patterns qualitatively correlated with SERR parameters.

© 2016 Elsevier LtdFlutter in aeronautical panels is a self-excited aeroelastic phenomenon which occurs during supersonic flights due to dynamic instability of inertia, elastic and aerodynamic forces of the system. In the flutter condition, when the critical aerodynamic pressure is reached, the vibration amplitudes of the panel become dynamically unstable and increase exponentially with time, significantly affecting the fatigue life of the existing aeronautical components. Thus, in this paper, the interest is to investigate the possibility reducing the effects of the supersonic aeroelastic instability of rectangular plates by applying passive constrained viscoelastic layers. The rationale for such study is the fact that as the addition of viscoelastic materials provides decreased vibration amplitudes it becomes important to quantify the suppression of plate flutter coalescence modes that can be obtained. Moreover, despite the fact that much research on the suppression of panel flutter has been carried out by using passive, semi-active and active control techniques, few works have been proposed to deal with the problem of predicting the flutter boundary of aeroviscoelastic systems, since they must conveniently account for the frequency- and temperature-dependent behavior of the viscoelastic material. After the presentation of the theoretical foundations of the methodology, the description of a numerical study on the flutter analysis of a three-layer sandwich plate is addressed.

© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.Installed jet noise is studied by means of a simplified configuration comprising flat rectangular plates situated in the nearfield of a round jet. Acoustic measurements are performed using a traversable 18-microphone azimuthal array, providing pressure data at 360 points on a cylindrical surface surrounding the jet-plate system. A rigid aluminium plate and a flexible, composite plate were tested to assess the influence of the plate stiffness on the scattered sound. The numerical predictions are confirmed by experiments and suggest that a reduction in the scattered sound level can be achieved as the flexibility of the plate is increased.

© 2016, European Conference on Composite Materials, ECCM. All rights reserved.Shape Memory Alloy Hybrid Composite (SMAHC) laminates are built with continuous carbon fibers and Shape Memory Alloy (SMA) wires, both embedded in a polymeric matrix thereby forming a three constituent composite material. The SMA actuation is triggered by temperature changes, resulting in modifications in the structural responses of SMAHC laminates. A particularly important structural characteristic of SMAHC laminates which is investigated in this paper is the aeroelastic stability boundary of flutter. The derivation of the effective mechanical properties of the SMAHC is based on micromechanical model which accounts for temperature and fraction of martensite/austenite transformation phases of the shape memory alloy. The mathematical problem is formulated using Hamilton's principle, allowing for derivation of the equilibrium equations and boundary conditions of the aeroelastic response. The governing equations are then discretized and solved by the finite element method. A parametric study is conducted where different geometric configurations, laminate stacking sequence, boundary conditions and curvatures are investigated. It is observed that the SMAHC structure is stabilized against flutter by proper tailoring of stiffening effects induced by the changes in the fraction of martensite/austenite transformation phases of the SMA. Therefore, it is possible to increase critical flutter speed by controlling the temperature of the SMA wires.

This work presents a successful methodology for obtaining failure envelopes of unidirectional fiber reinforced composites based on micromechanical analysis by the asymptotic homogenization [1] method. Given a structure (in this case a composite material lamina) and external loads, plus having its material heterogeneity geometrically represented by a periodic unit cell of microstructure, the asymptotic homogenization method is able to predict its micromechanical stresses. Such stresses may be evaluated by failure criteria of the composite's constituents for several loading conditions, and this way it is possible to assess the composite's failure envelope. In the methodology developed, a periodic unit cell of the composite microstructure is isolated, consisting of a parallelepiped of polymeric matrix reinforced by cilindrical fibers oriented in one direction, and its behaviour is evaluated by an appropriate finite elements model. At first, the unit cell is tested in several directions to find strengths for the matrix and fiber and also matrix/fiber interface, thus evaluating failure characteristics of the composite constituents. The tests are carried out considering several possible orientations for the unit cell inside the related macroscopic media, which take into account the possible relative positions of the chosen unit cell inside the material heterogeneity of the composite. Then, the strengths of the constituents are used to predict the failure envelope for the fiber reinforced material, according to failure criteria devoted to the constituents. In this prediction, it is possible to say which is the constituent that fails first for each of the test loads. The results obtained are in good agreement with experimental data for carbon/epoxi and glass/epoxi composites. Moreover, the envelopes obtained are similar to the Puck & Schürmann [2] criterion, widely used to predict failure of such composites. This way, the present methodology renders good failure envelopes for fiber reinforced composites and gives information on the strengths of the constituents and material phase of failure, benefits from an incorporated micromechanical analysis.

© 2015 Elsevier Ltd. All rights reserved.Carbon fiber reinforced polymers (CFRP) structure can include dropping-off plies in order to comply with design requirements aiming at significant weight savings. However this type of discontinuity represents a potential source of delamination initiation and propagation which requires assessment of the mechanisms acting at the crack tip. This research investigates the influence of delamination modes I and II on the overall damage process observed in CLS specimen subjected to cyclic loads. The main contribution of this work focuses on the identification and physical interpretation of complex failure mechanisms in harness satin fabric. For this purpose a detailed fractographic analysis was carried out to qualitatively assess the surface fractures in these type of laminates. Results obtained for cyclic loaded CLS specimens were compared to analytical closed form solutions available in the literature. Results indicated that delamination front exhibited distinguishable delamination modes I and II propagating at constant mixed mode ratio (G<inf>I</inf>/G<inf>T</inf>).

© SAGE Publications.Currently, the standard delamination tests established by ASTM (available for mode I and mix mode) are limited to unidirectional composites. Although some researchers have conducted delamination tests in woven composites, their information is still limited. In order to understand the propagation behavior and the value of fracture toughness of woven composites, this article evaluated a 5HS carbon/epoxy composite with a weft-dominated surface. Tests were conducted in double cantilever beam and end notch flexure configurations using an energy-based approach for data reduction in modes I and II, respectively. The results were assessed in terms of delamination resistance curves (R-curves). Both delamination modes showed consistent behaviors for extending the application of the standard procedures, as the energy variation can describe well the crack growth dependence of the irregular surface caused by the crimp, which was more pronounced for mode I.

© 2015 Elsevier Masson SAS. All rights reserved.This paper presents a detailed investigation on the post-buckling behaviour of adhesively bonded stiffened panels subjected to in-plane shear loading. An experimental programme was carried to determine the buckling load, buckling shape, collapse load and failure modes of two bonded stiffened panels. A nonlinear finite element based modelling approach, accounting for geometrical and material nonlinearities as well as progressive failure in the adhesively bonded interface between the skin and the stiffener is proposed to predict the structural behaviour of the panels up to failure. This approach consists in modelling the bonded interfaces using a newly developed cohesive zone based constitutive damage model. In order to account for damage in the stiffener and the skin a Von Mises based constitutive damage model is also formulated and presented in the paper. Both constitutive models were implemented into ABAQUS/Explicit finite element code as user-defined material models. A very good agreement between experimental results and numerical predictions is obtained using the proposed modelling approach, with deviations smaller than 8% in buckling load and bonded interface failure load onset.

© 2014 SAGE Publications.This paper presents a continuum damage mechanics failure model to predict mixed-mode delamination growth in composite laminates subjected to static and high-cycle fatigue loading. The proposed formulation has been developed for robust nonlinear finite element formulations based on explicit direct time integration schemes, particularly the central difference method. The failure model has been implemented as a user-defined material model into ABAQUS/Explicit finite element code within C3D8 hexahedron solid elements. Numerical simulations were performed at coupon level for double cantilever beam, end-notched-flexure, mixed-mode bending and mixed-mode flexure specimens. Predictions obtained using the proposed failure model were compared with experimental results available in the open literature. Good agreement between numerical and experimental results was found.

© 2014 Elsevier Ltd.The small mass impact, such as runway debris and dropped tools, is a major issue during the design of composite plates, as it may significantly reduce the plate strength and stiffness without any visible damage. This work presents a novel optimization procedure to maximize the critical buckling load of composite plates subjected to small mass impact considering impact response and delamination threshold load predicted by closed forms solutions. The optimization is based on lamination parameters combined with laminate databases. In the first phase of the optimization a simplified model is used to obtain an approximate optimum. A design optimization using refined finite element model starting from the previous phase optimum is performed. The proposed algorithm is efficient, robust and applicable to many optimization problem for composite plates.

© 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All Rights Reserved.Trailing edge scattering is a significant source of sound in aeroacoustics, and elasticity is known to decrease the radiated sound by a process involving coupled acoustic and bending waves. Most of the analysis in the literature is appropriate for metallic plates, which are isotropic. We extend a numerical method, based on the solution of a boundary element method (BEM) with boundary conditions given by the structural problem, to account for anisotropic, composite plates. We perform a comparison between composite and metallic plates with the same thickness and similar bending stiffness. For both cases, elasticity is seen to reduce the scattered sound; composite plates lead to greater reductions of far-field sound due to their lower specific mass and consequent higher fluid loading factor. Results also show that orientation of laminae can be used so as to optimise plates for acoustic radiation at specific Helmholtz numbers k0 of interest: different lay-ups present changes in structural resonance frequencies, and higher acoustic benefits can be obtained by ensuring that a given k0 is between two resonances, in a situation where acoustic excitation and elastic response are in phase opposition.

Emergence of flutter compromises not only the long term durability of the wing structure, but also the operational safety, flight performance and energy efficiency of the aircraft. Effective means of flutter prevention are, therefore, mandatory in the certification of new flight vehicles. This work intends to address the application of viscoelastic material for flutter suppression in a typical section under quasi-steady and unsteady aerodynamic loads. A numerical procedure is proposed to solve the set of non-linear equations. A parametric study showing the influence of the temperature on the stiffness, damping and flutter velocity of the system is also carried out using different viscoelastic based damping arrangements. The preliminary results indicate that the aeroelastic behavior of the system is significantly affected by the temperature and viscoelastic damping arrangement.

Nondestructive experimental methods to calculate the buckling load of imperfection sensitive thin-walled structures are one of the most important techniques for the validation of new structures and numerical models of large scale aerospace structures. Vibration correlation technique (VCT) allows determining equivalent boundary conditions and buckling load for several types of structures without reaching the instability point. VCT is already widely used for beam structures, but the technique is still under development for thin-walled plates and shells. This paper intends to explain the capabilities and current limitations of this technique applied to two types of structures under buckling conditions: flat plates and cylindrical shells prone to buckling. Experimental results for a flat plate and a cylindrical shell are presented together with reliable finite element models for both cases. Preliminary results showed that the VCT can be used to determine the realistic boundary conditions of a given test setup, providing valuable data for the estimation of the buckling load by finite element models. Also numerical results herein presented show that VCT can be used as a nondestructive tool to estimate the buckling load of unstiffened cylindrical shells. Experimental tests are currently under development to further validate the approach proposed herein. © 2014 Elsevier Ltd.

© 2014 Elsevier Ltd. All rights reserved.This chapter presents a comprehensive review on existing approaches for modeling damage in composite structures. Special emphasis is given in robust and reliable constitutive models that enable prediction of failure initiation and failure progression in composite laminates within an unified way. Theoretical and numerical issues related to intra and interlaminar failure modeling are also discussed in detail. The constitutive models formulations presented in this chapter are based on the Continuum Damage Mechanics (CDM) approach and enables the control of the energy dissipation associated with each failure mode regardless of mesh refinement and fracture plane orientation. Within the CDM context, internal thermodynamically irreversible damage variables are defined in order to quantify damage concentration associated with each possible failure mode enabling the prediction of the gradual stiffness reduction for each composite ply. Numerical examples are also provided in order to illustrate the models capabilities.

Interlaminar fracture process in woven composite laminate in general takes place in a nonlinear front propagation when analyzing a 2-D propagation for mode I opening. The nonlinearity is justified by the arresting crack at the warp yarn. Regarding mode II opening, the crack interaction to the crimp surface still needs further assessment. The tortuous crack path and, consequently, loading directions on both modes were determined in this work by fractographic analysis of quasi-static and cyclic loading fractured in 5HS carbon/epoxy composite specimens. The crack growth rate and strain energy release rate were compared to the fractographic images that approximately pointed out the driving force in each delamination stage. Mode I fracture surface showed to have more influence on fracture toughness (GIc) than amplitude loading while mode II fracture surface suggested the opposite behavior, all based on the well-known fracture patterns established along the years for other composite architectures.

© 2014, Journal of Aerospace Technology and Management. All Rights Reserved.The developments of innovative adaptive structures on Unmanned Aerial Vehicles (UAVs), such as morphing wings, can potentially reduce system complexities by eliminating control surfaces and their auxiliary equipment. This technology has the potential of allowing a UAV to adapt to different mission requirements or to execute a particular mission more effectively by maintaining an optimum airfoil section over a range of speeds for different segments of a mission profle. Studies on a number of smart materials candidates are currently available in the open literature to achieve wing morphing. The material selection depends on several factors including fast dynamic response, low weight, capability to operate over a wide range of fight conditions and low power consumption. This paper presents a review on smart materials technologies for UAV morphing wings. A numerical study in terms of power requirements is also presented for two morphing wing concepts: fapped and twisted wing planforms. The energy calculations for both morphing confgurations were based on a two-step procedure. The frst step consists of computing the aerodynamic energy using an in-house Vortex-Lattice (VL) based program. Subsequently the pressure feld obtained from the frst step is then mapped into a fnite element mesh and the structural strain energy is calculated. The numerical results indicated that fapped morphing wings have a better aerodynamic performance when compared to twisted wings and different morphing levels can be achieved using lighter smart materials with lower specifc energy for this confguration.

This chapter presents a comprehensive review on existing approaches for modeling damage in composite structures. Special emphasis is given in robust and reliable constitutive models that enable prediction of failure initiation and failure progression in composite laminates within an unified way. Theoretical and numerical issues related to intra and interlaminar failure modeling are also discussed in detail. The constitutive models formulations presented in this chapter are based on the Continuum Damage Mechanics (CDM) approach and enables the control of the energy dissipation associated with each failure mode regardless of mesh refinement and fracture plane orientation. Within the CDM context, internal thermodynamically irreversible damage variables are defined in order to quantify damage concentration associated with each possible failure mode enabling the prediction of the gradual stiffness reduction for each composite ply. Numerical examples are also provided in order to illustrate the models capabilities. © 2014 Elsevier Ltd All rights reserved.

This work investigates the numerical prediction of compression after impact strength in woven composite laminates. Intralaminar and interlaminar damage prediction were evaluated using proposed damage models implemented as user defined material in ABAQUS Explicit multipurpose FE code. The numerical models were developed using the finite elements method with two different modeling approaches named Single Shell Model (SSM) and Split Shell Model (SpSM). The Single Shell Model (SSM) used only shell elements to model the laminates and the delamination effects were neglected. The delamination effects were included in the Split Shell Model (SpSM) by using a delamination contact-logic. An experimental programme was carried out to validate the proposed damage modeling approaches. The proposed damage models and the modeling approaches have proven to be capable of reproducing experimental results with good accuracy for the impact tests and CAI tests. © 2014 Elsevier Ltd.

Fractography involves a detailed study of the fractured surface morphology of materials. Usually the Scanning Electron Microscopy (SEM) technique is applied to investigate the cause of failure and the relationship between failure modes and the microstructure of the material under investigation. This information allows one to relate structure, processing methods and materials properties with the onset in their failure and propagation. In this study, fractographical analysis is employed to investigate the failure aspects induced by Mode II delamination of structural fiber-reinforced polymer composite laminates. Samples with Teflon® inserted at the mid-plane of one end of the laminate (end-notched specimens) were subjected to four-point bending (4-ENF) test. The samples were prepared from laminates manufactured in autoclave with prepreg layers (0°, 90°) of IM7 plain weave woven fabric fiber architecture and M21-epoxy resin. The results indicate that the fracture surface is very irregular where fibers and fiber prints at the matrix are visualized, along with fractographic aspects named cusps and scallops, formed during the polymeric matrix shearing.

The wind energy research has grown substantially in the past few years, considerably fostered by the pursuit for a clean and sustainable energy source. Improvements on the design methods are increasingly needed. The purpose of this research is to investigate the use of the Loewy's lift deficiency function (LDF), also named Returning Wake Model, coupled with a non-stationary Blade Element-Momentum Method (BEM). The LDF simulates the influence of the wake behind the wind turbine on its capacity to generate power. It is expected that this model reduce the dependency of the several empirical parameters necessary in other wake models which are currently used. Aiming to validate the results obtained in this new approach they are compared with those provided by commercial computational software and they have proven to be very consistent. It is concluded that the method is feasible to be used as an efficient design and optimization tool of upwind horizontal axis wind turbine blades.

This work investigates the aeroelastic stability boundary of flutter in aircraft composite panels, curved or flat, subject to the effect of stress stiffening caused by the piezoelectric actuator (PZT). Hamilton's principle is used for the formulation of the energy functional and to obtain the equilibrium equations and boundary conditions of the problem. The finite element method is employed to numerically solve the equations. The aeroelastic behavior of panels manufactured in composite material (boron-epoxy) or conventional material (aluminum 2024-T3) are assessed. Two layers of piezoelectric material (ACX QP10N) are attached to the panels: one on the top surface one on the bottom surface of the panels. Prescribed voltages are statically applied to the piezoelectric actuators, inducing a prestress field which is responsible for the stress stiffening effects when coupled with the nonlinear strain components. Different geometric configuration, laminate stacking sequence, boundary conditions and curvatures are investigated. The study shows that mechanically strain-induced piezoelectric effect increases the rate of occurrence of flutter, stabilizing the plate. This stiffening of the structure is related to the voltage applied on the actuators and the geometrical parameters of the plate. Thus, one can control the occurrence of flutter speed by controlling the voltage applied and the proper design of the geometric properties of the panel and tailoring of the composite laminate. © 2012 Elsevier Ltd.

Many components of modern aircrafts are now manufactured from polymer composites. Reinforced laminates with continuous carbon fibers and modified epoxy resin are employed in primary and secondary structures to reduce weight and improve the aircraft performance. However, if a circumstantial failure happens, the complex fracture process of the laminates may involve interlaminar damage mechanisms. The delamination is the interlaminar discontinuity which may propagate catastrophically with the application of mechanical loads. The Double Cantilever Beam (DCB) is the most used method to determine the Mode I fracture toughness of structural composites. In this work samples prepared from a plain weave fabric laminate were submitted to Mode I delamination under static load at room temperature. The analysis of the delaminated surfaces was performed with scanning electron microscopy (SEM). The results show that the fracture process initiates at the resin pockets after a Teflon® insert and propagates along the resin rich areas at the crossing of weft and warp tows. The main fractographical aspects revealed are identified, reported and discussed.

This paper presents a ballistic impact simulation of an armour-piercing projectile in hybrid ceramic/fiber reinforced composite armour. The armour is composed by an alumina plate and an ultra high molecular weight polyethylene composite. In order to model the armour behavior three different constitutive models were formulated and implemented into ABAQUS/Explicit finite element code. Comparisons between numerical predictions and experimental results in terms of damage shape/extent and V 50 are also presented and discussed in the paper. © 2011 Elsevier Ltd. All rights reserved.

The nonlinear response of initially imperfect composite plates with piezoelectric actuators is investigated. The nonlinearity is limited to the prebuckling regime, where higher order terms present in the strain energy expression can be neglected. The advantage of the electromechanical coupling is exploited in two ways. First, the in-plane piezoelectric stress stiffening effect is used to tailor a stress distribution that inherently increases the critical buckling loads of perfect composite plates by posing an optimization problem that efficiently handles eventual uncertainties involved in the application of mechanical loadings. Second, piezoelectric bending moments are applied in order to avoid or ameliorate the undesirable effects of initial imperfections. An actuation strategy, where the piezoelectric membrane forces and bending moments are decomposed via an appropriate selection of voltages applied to piezoelectric patches that are symmetrically bonded to the top and bottom surfaces of the plate, is proposed and shown to be effective. © 2011 American Society of Mechanical Engineers.

This paper presents a detailed experimental and numerical investigation on the structural behavior of stiffened composite panels subjected to in-plane shear loads. The experimental work includes the development of a test device for post-buckling analyses of laminated panels subjected to shear loads. The panels out-of-plane displacement field in the post-buckling regime was experimentally characterized using a non-contact 3-D optical device. A test procedure was proposed to obtain reliable and reproducible results. The following parameters were established: geometry and instrumentation of the specimens, test mechanisms, data acquisition procedures and analysis procedures for test data.The numerical objective of this work is to implement a modeling methodology for analysis of composite stiffened panels using finite elements. The proposed methodology takes into account large displacements and material nonlinearity effects by using a damage mechanics based progressive failure model.Preliminary results for tested specimens with the proposed configuration indicate that the stiffened composite shear webs have significant post-buckling strength. © 2010 Elsevier Ltd.

This paper presents a numerical study on the impact resistance of composite shells laminates using an energy based failure model. The damage model formulation is based on a methodology that combines stress based, continuum damage mechanics (CDM) and fracture mechanics approaches within a unified procedure by using a smeared cracking formulation. The damage model has been implemented as a user-defined material model in ABAQUS FE code within shell elements. Experimental results obtained from previous works were used to validate the damage model. Finite element models were developed in order to investigate the pressure and curvature effects on the impact response of laminated composite shells. © 2010 Elsevier Ltd.

This paper presents a detailed numerical investigation of the post-buckling behavior of composite shear webs using the finite element method. The numerical analysis accounts for material and geometric non-linearity effects and has been divided into three steps. The first step consists of computing the critical buckling loads as well as their corresponding buckling modes. Geometric imperfections described approximately in terms of linear combinations of different normal modes are then introduced into the model. Finally a quasi-static analysis is carried out including a progressive failure model. The progressive failure model has been implemented as a user defined material model within shell elements in Abaqus/Explicit finite element code. © Taylor & Francis Group, LLC.

A technique for enhancement of buckling loads of composite plates is proposed. The technique relies on using stress stiffening to create a non-zero tensile force acting along the plate plane which ultimately permits the application of higher external compressive forces that lead to traditional buckling instabilities. The idea is to completely restrain the plate movements in its plane direction, at all edges, and to apply voltages to pairs of symmetrically bonded piezoelectric patches. This voltage is applied such that the piezoelectric patches contract resulting in a uniform tensile force over the plate plane.

This work deals with a numerical investigation on the structural behaviour of stiffened composite panels subjected to in-plane shear loads in the post-buckling regime. The modelling approach takes into account large deformations and material nonlinearity effects by using a damage mechanics based progressive failure model.

This paper presents a 3-D failure model for predicting the dynamic material response of composite laminates under impact loading. The formulation is based on the Continuum Damage Mechanics (CDM) approach and enables the control of the energy dissipation associated with each failure mode regardless of mesh refinement and fracture plane orientation. Internal thermodynamically irreversible damage variables were defined in order to quantify damage concentration associated with each possible failure mode and predict the gradual stiffness reduction during the impact damage process. The material model has been implemented into LS-DYNA explicit finite element code within solid elements and it has proven to be capable of reproducing experimental results with good accuracy in terms of static/dynamic responses, absorbed energy and extent of damage. © 2007 Elsevier Ltd. All rights reserved.

This paper presents an analytical model for the prediction of the elastic behaviour of plain-weave fabric composites. The fabric is a hybrid plain-weave with different materials and undulations in the warp and weft directions. The derivation of the effective material properties is based on classical laminate theory (CLT). The theoretical predictions have been compared with experimental results and predictions using alternative models available in the literature. Composite laminates were manufactured using the resin infusion under flexible tooling (RIFT) process and tested under tension and in-plane shear loading to validate the model. A good correlation between theoretical and experimental results for the prediction of in-plane properties was obtained. The limitations of the existing theoretical models based on classical laminate theory (CLT) for predicting the out-of-plane mechanical properties are presented and discussed. © 2007 Elsevier Ltd. All rights reserved.

A numerical and experimental investigation on the mode-I intralaminar toughness of a hybrid plain weave composite laminate manufactured using resin infusion under flexible tooling (RIFT) process is presented in this paper. The pre-cracked geometries consisted of overheight compact tension (OCT), double edge notch (DEN) and centrally cracked four-point-bending (4PBT) test specimens. The position as well as the strain field ahead of the crack tip during the loading stage was determined using a digital speckle photogrammetry system. The limitation on the applicability of the standard data reduction schemes for the determination of intralaminar toughness of composite materials is presented and discussed. A methodology based on the numerical evaluation of the strain energy release rate using the J-integral method is proposed to derive new geometric correction functions for the determination of the stress intensity factor for composites. The method accounts for material anisotropy and finite specimen dimension effects regardless of the geometry. The approach has been validated for alternative non-standard specimen geometries. A comparison between different methods currently available for computing the intralaminar fracture toughness in composite laminates is presented and a good agreement between numerical and experimental results using the proposed methodology was obtained. © 2006 Elsevier Ltd. All rights reserved.

This paper presents an experimental and numerical study focused on the tensile fibre fracture toughness characterisation of hybrid plain weave composite laminates using non-standardized Overheight Compact Tension (OCT) specimens. The position as well as the strain field ahead of the crack tip in the specimens was determined using a digital speckle photogrammetry system. The limitation on the applicability of standard data reduction schemes for the determination of the intralaminar fibre fracture toughness of composites is presented and discussed. A methodology based on the numerical evaluation of the strain energy release rate using the J-integral method is proposed to derive new geometric correction functions for the determination of stress intensity factor for alternative composite specimen geometries. A comparison between different methods currently available to compute the intralaminar fracture toughness in composites is also presented and discussed. Good agreement between numerical and experimental results using the proposed methodology was obtained.

A criterion is derived for delamination onset in transversely isotropic laminated plates under small mass, high velocity impact. The resulting delamination threshold load is about 21% higher than the corresponding quasi-static threshold load. A closed form approximation for the peak impact load is then used to predict the delamination threshold velocity. The theory is validated for a range of test cases by comparison with 3D finite element simulation using LS-DYNA and a newly developed interface element to model delamination onset and growth. The predicted delamination threshold loads and velocities are in very good agreement with the finite element simulations. Good agreement is also shown in a comparison with published experimental results. In contrast to quasi-static impacts, delamination growth occurs under a rapidly decreasing load. Inclusion of finite thickness effects and a proper description of the contact stiffness are found to be vital for accurate prediction of the delamination threshold velocity. © 2005 Elsevier Ltd. All rights reserved.

Piezoelectric actuators are usually mounted to the top and bottom surfaces of plates and may induce in-plane extension, bending and localized shear deformations at the structural element. The in-plane stresses may have a significant influence on the mechanical behavior of thin plates as initial and/or residual stresses affect the flexural stiffness and in turn the dynamic and stability characteristics of plates. In this work, the effect of the in-plane piezoelectric induced stresses on the natural frequencies of composite plates is numerically and experimentally investigated. A finite element formulation is presented for the analysis of laminated plates with an arbitrary number of piezoelectric actuators and/or sensors. Von Kàrmàn non-linear strain-displacement relations are used and ideal linear behavior is assumed for the piezoelectric actuation. The problem is decomposed into an in-plane problem where the strain field induced by the piezoelectric actuators is computed. The natural frequencies and vibration modes are then computed taking the stress stiffening effects of these piezoelectric stresses into account. A number of different configurations are numerically and experimentally analyzed to verify the proposed theory. The configurations use eight PZT actuators bonded to three layer glass fiber/epoxy plates. The plates are square and clamped along two opposing edges and free along the other two. Good agreement is obtained between the predicted and measured natural frequencies. © 2002 Elsevier Science Ltd. All rights reserved.