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In this work, a preliminary numerical assessment on the application of an additive manufactured hybrid metal/composite shock absorber panels to a military seat ejection system, has been carried out. The innovative character of the shock absorber concept investigated is that the absorbing system has a thickness of only 6 mm and is composed of a pyramid-shaped lattice core that, due to its small size, can only be achieved by additive manufacturing. The mechanical behaviour of these shock absorber panels has been examined by measuring their ability to absorb and dissipate the energy generated during the ejection phase into plastic deformations, thus reducing the loads acting on pilots. In this paper the effectiveness of a system composed of five hybrid shock absorbers, with very thin thickness in order to be easily integrated between the seat and the aircraft floor, has been numerically studied by assessing their ability to absorb the energy generated during the primary ejection phase. To accomplish this, a numerical simulation of the explosion has been performed and the energy absorbed by the shock-absorbing mechanism has been assessed. The performed analysis demonstrated that the panels can absorb more than 60% of the energy generated during the explosion event while increasing the total mass of the pilot-seat system by just 0.8%.

Composites design has always been perceived to be complicated, time-consuming and therefore not affordable. With a paradigm shift, it is possible to make composites are as straightforward as metals. Each metal is characterized by two key material properties: its Young’s modulus and strength. The adoption of the Tsai’s modulus as the only stiffness constant characterising any CFRP, and failure strain as the only measure of strength, we can make composite design as straightforward as metals design. Actually, with these two material characteristics, for any CFRP, material, laminate, and strain allowable can all be scaled, just like metals. Another break-through is the replacement of the traditional quad laminates of 0, ±45, 90 by double-double (DD) in [±Φ/±ψ]. This replacement allows to substuture a discrete collection by a continuous field of orientations. A few dozen choices, replaced by hundreds. Indeed, double-double can make structures lighter that quad cannot do, thanks to the Lam-search optimizer, which is able, in combination with Fem tools, to sort, instantaneously, the best DD angles allowing laminates stiffness and strength tailored on specific loading conditions. The bottom line: Tsai’s modulus and double-double are new but are simple conceptually, and Lam-search makes their use practical. In the end, lower weight and lower cost are derived by DD to deliver aggressive taper and 1-axis layup that will be stiff, strong, easy to layup, not prone to error, wrinkle, warpage and delaminatnion.

In this work, an experimental characterization of Shape Memory Alloy (SMA) springs was carried-out to assess the efficiency of simple analytical procedure for a preliminary design of Shape Memory Alloy-based spring actuators. Two static analytical models were considered recursively, respectively to determine SMA material properties by accurate tensile experimental test and to identify the optimized geometrical characteristic of SMA-bias spring-based unit for a specific feasible mini-actuator for vehicle moveable part. The final extra-force and displacement of the actuator was validated at different temperature above transformation temperature to assess the limitation of the implemented procedure.

The capability to ensure passenger safety is a core requirement of any transportation facility. For this reason, the concept of Crashworthiness, i.e. the capacity offered by a structure to absorb shocks during an accident and preserve occupant safety, over the years, gained a major role in the designing of transportation devices. Indeed, considerable effort has been expended by researchers in the study of shock-absorbing devices able to increase the safety performance. Actually, this paper introduces both a numerical and experimental assessment of the effectiveness of a new hybrid (metal-composite) shock absorber concept, fabricated by means of Additive Manufacturing. Additive technology allows some fundamental benefits such as the possibility to manufacture cores with a complex shape in sandwich panels , able to maximize the energy absorption capabilities. A numerical investigation on the energy absorption capabilities of different geometrical and material configurations has been carried out, to select the best performing configurations to be manufactured and tested to impact. Impact tests on core, skins and complete sandwich panels have been carried out to assess the capability to absorb impact energy under different boundary conditions.

A preliminary design of customized antibiotic-loaded poly-methyl-methacrylate (ALPMMA) spacer characterized by an appropriate footprint according to the specific patient’s anatomy and a reliable mechanical response to severe functional loads (i.e., level walking and 45° bent knee) is reported. The targeted virtual prototyping process takes origin from a novel patented 3D geometrical conceptualization characterized by added customization features and it is validated by a preliminary FEM-based analysis. Mechanical and thermomechanical properties of the antibiotic-doped orthopedic PMMA cement, which will be used for the future prototype manufacturing, were measured experimentally by testing samples taken during a real day-running orthopedic surgery and manufactured according to the surgeon protocol. FEM analysis results indicate that small area is subjected to intensive stresses, validating the proposed geometry from the mechanical point of view, under the two loading scenarios, moreover the value of safety margins results positive, and this is representative of the lower stress magnitude compared to the critical material limits. The experimental data confirm that the presence of antibiotic will last during the surgeon period moreover, the temperature dependent modulus of the bone cement is slightly affected by the body range temperature whereas it will drastically drop for higher temperature out the range of interest. A complete customization, according to a patient anatomy, and the corresponding real prototype spacer will be manufactured by 3D printing techniques, and it will be validated by destructive testing during the second stage of this activity before commercialization.

Shape memory alloys (SMAs) belong to a class of smart materials with the ability to memorize or retain their previous shape when subjected to external stimuli such as thermomechanical or magnetic gradients. SMAs have attracted significant attention and interest in in a broad range of commercial applications owing to their unique and superior properties, which have been well-analyzed and characterized by applied research studies. The key benefit of SMA actuators lies mainly in the possibility of simple and compact mechanical designs requiring low actuating force levels. This chapter provides a timely review of SMA research and commercial applications with special relevance to the automotive technology field. Advances and application opportunities are discussed that make them ideally suited for industrial challenges.

This work provides a feasibility and effectiveness analysis, through numerical investigation, of metal replacement of primary components with composite material for an executive aircraft wing. In particular, benefits and disadvantages of replacing metal, usually adopted to manufacture this structural component, with composite material are explored. To accomplish this task, a detailed FEM numerical model of the composite aircraft wing was deployed by taking into account process constraints related to Liquid Resin Infusion, which was selected as the preferred manufacturing technique to fabricate the wing. We obtained a geometric and material layup definition for the CFRP components of the wing, which demonstrated that the replacement of the metal elements with composite materials did not affect the structural performance and can guarantee a substantial advantage for the structure in terms of weight reduction when compared to the equivalent metallic configuration, even for existing executive wing configurations.

Composite materials are characterized by extraordinary strength and stiffness properties which can be strongly affected by intralaminar and interlaminar failure modes. Among the different damages, delaminations can be considerably influenced by the fatigue induced material properties degradation. Indeed, delamination propagation phenomena could begin at a fairly low number of cycles, progressively leading to the growth of the delaminated area and to the decrease of the global structural stiffness until the structural catastrophic failure. In this paper, a numerical procedure for fatigue induced delamination simulation has been developed. The mechanical behaviour of composite panels under cyclic loading conditions have been numerically investigated. The proposed approach, implemented in ANSYS® FEM software by using the Ansys Parametric Design Language (APDL) adopts a mesh and time step independent procedure based on the Virtual Crack Closure Technique (VCCT), combined with the Paris-Law formulation, which permits to take into account the local damage accumulation along the delamination front.

Thanks to their peculiar morphing capabilities, Shape Memory Alloys (SMA) materials are being increasingly adopted in many engineering fields, such as automotive and aerospace. Indeed, their pseudo-elastic behavior and their shape memory effect, which allows the recovery of an initial state after a deformation, make them particularly suitable for switch actuators and morphing structures applications. In this work, the behavior of SMAs is investigated. Then, a user defined material model (UMAT) has been implemented in the Abaqus Standard Finite Element (FE) environment to numerically simulate the behavior of SMAs, including their pseudo-elastic behavior and shape memory effect. The material model, validated by means of comparisons with analytical data, has been used to investigate the feasibility of a bi-stable biased actuator operated by SMA springs.

This paper aims to investigate the crashworthiness capability of a commercial aircraft metallic sandwich leading edge, subjected to bird strike events. A sensitivity analysis is presented, aimed to assess the influence of the skin parameters (inner and outer faces and core thicknesses) on the leading-edge crashworthiness and to determine, among the configurations able to withstand a bird strike event, the best compromise in terms of weight and structural performances. In order to easily manage the design parameters and the output data, the ModeFrontier code was used in conjunction with the FE code Abaqus/Explicit. A dedicated python routine was developed to define a fully parametric simplified leading-edge model. To fulfill the aerodynamic requirements, the external surfaces were considered fixed during the sensitivity analysis, and, thus, only the internal leading edge&rsquo, s components were modified to study their influence on the structural response. The total mass of the model, the maximum deformation and the energy dissipated due to material failure and the plastic deformations were monitored and used to compare and assess the behavior of each configuration.

This work is focused on the investigation of the structural behavior of a composite floor beam, located in the cargo zone of a civil aircraft, subjected to cyclical low-frequency compressive loads with different amplitudes. In the first stage, the numerical models able to correctly simulate the investigated phenomenon have been defined. Different analyses have been performed, aimed to an exhaustive evaluation of the structural behavior of the test article. In particular, implicit and explicit analyses have been considered to preliminary assess the capabilities of the numerical model. Then, explicit non-linear analyses under time-dependent loads have been considered, to predict the behavior of the composite structure under cyclic loading conditions. According to the present investigation, low-frequency cyclic loads with peak values lower than the static buckling load value are not capable of triggering significant instability.

The thermo-structural design of the wing leading edge of hypersonic vehicles is a very challenging task as high gradients in thermal field, and hence high thermal stresses, are expected. Indeed, when employing passive hot structures based thermal protection systems, very high temperatures (e.g., 1400 &deg, C) are expected on the external surface of the wing leading edge, while the internal structural components are required to not exceed a few hundred degrees Celsius (e.g., 400 &deg, C) at the interface with the internal cold structure. Hence, ceramic matrix composites (CMC) are usually adopted for the manufacturing of the external surface of the wing leading edge since they are characterized by good mechanical properties at very high temperatures (up to 1900 &deg, C) together with an excellent thermal shock resistance. Furthermore, the orthotropic behavior of these materials together with the possibility to tailor their lamination sequence to minimize the heat transferred to internal components, make them very attractive for hot structure based thermal protection systems applications. However, the numerical predictions of the thermo-mechanical behavior of such materials, taking into account the influence of each ply (whose thickness generally ranges between 0.2 and 0.3 mm), can be very expensive from a computational point of view. To overcome this limitation, usually, sub-models are adopted, able to focus on specific and critical areas of the structure where very detailed thermo-mechanical analyses can be performed without significantly affecting the computational efficiency of the global model. In the present work, sub-modeling numerical approaches have been adopted for the analysis of the thermo-mechanical behavior of a ceramic matrix composite wing leading edge of a hypersonic vehicle. The main aim is to investigate the feasibility, in terms of computational efficiency and accuracy of results, in using sub-models for dimensioning complex ceramic matrix components. Hence, a comprehensive study on the size of sub-models and on the choice of their boundaries has been carried out in order to assess the advantages and the limitations in approximating the thermo-mechanical behavior of the investigated global ceramic matrix composite component.

In the present work, a numerical study on the dynamic response of a composite fuselage barrel, in relation to crashworthiness, has been investigated. The aim of this work is to investigate the influence of the material fracture toughness on the capability of a composite fuselage barrel to tolerate an impact on a rigid surface. Three different material configurations with different intra-laminar fracture energy values were considered to take into account variations in material toughness. Indeed, the dynamic behaviour of the analysed fuselage barrel has been numerically simulated by means of the FE (Finite Element) code Abaqus/Explicit. The effects of intralaminar fracture energy variations on the impact deformation of the barrel has been evaluated comparing the numerical results in terms of displacements and damage evolution for the three analysed material configurations.

In this work, the mechanical behaviour of natural composite fibre sandwich panels is experimentally investigated. The composite sandwich panels are composed of a natural core made of Posidonia dried leaves reinforced with aluminium skins. The introduced material system is characterized by good structural behaviour (due to the aluminium skins) combined with the good thermal and acoustical insulation provided by the Posidonia core. Flexural and impact tests, characterized by different impact energies, have been performed on specimens with and without the aluminium skins, to preliminary assess the influence of the Posidonia core densities on the natural fibre sandwich mechanical behaviour.

Thanks to their excellent structural performance at high temperature, the Ceramic‐Matrix Composites (CMCs) are good candidates for light‐weight structures in high temperature applications, such as applications to re-entry vehicles. Beyond the well-known high temperature resistance, the CMCs do also show good damage tolerance properties, making them suitable for applications to large primary structures. Unfortunately, the production at the industrial level is currently at an embryonic stage due to the criticalities in the involved manufacturing processes which do not allow to produce defects free components. During the design phase, the presence of these manufacturing related defects are taken into account by introducing knock-down factors reducing the mechanical properties. However, this approach can be highly conservative and can strongly limit the intrinsic advantages of using such materials. Moreover, even if suitable non-destructive techniques (such as tomography) can be adopted to identify the manufacturing defects, still uncertainties exist about these defects evolution and, above all, about their influence on the structural performance (in term of stiffness and strength) of real components. The adoption of suitable material and fracture mechanics numerical models, could help to correctly predict the stiffness and strength characteristics of CMCs at coupon and subcomponent level. Which such advanced numerical models the manufacturing defects and their influence on the structural performance could be taken into account leading to a strong reduction of the currently adopted knockdown factors. The present work aims to investigate the damage tolerance of a C/C-SiC hot structure acting as aerodynamic control surface of a reusable re-entry vehicle. These vehicles are exposed to severe environmental conditions when re-entering Earth atmosphere. Indeed, the highest loaded areas such as nose, leading edges and control surfaces can experience temperatures up to 1650°C. Since the structural integrity of a re-entry vehicle needs to be guaranteed during all the mission phases and for multiple missions, a damage tolerant reliable design is required. Starting from the validation at coupon level of the material and fracture mechanics numerical models, by means of experimental data available in literature, a full parametric Finite Element model has been developed. The proposed parametric model is able to consider any delamination in any planar and thickness location all along the body flap domain. In order to reduce the computational costs, the entire body flap was modelled by means of layered shell elements, while layered solid elements have been used in the delaminated area. Hence, to connect the global coarse model and the local refined model, a global-local approach has been implemented. Finally, a sensitivity analysis has been performed finalised to assess the influence of location (in plane and trough the thickness) and dimension (radius) of a circular delamination on the damage tolerance of the investigated structure.

The aim of this work is to investigate the effects of thermal loads on the elastic instability of jointed composite components. Fibre-reinforced composite materials, which are of great interest for aerospace structures, are characterised by anisotropic thermal expansion coefficients. This characteristic influences the elastic stability behaviour when thermal stresses applied on the component. Actually, the presence of thermal coefficients varying with direction, can generate interlaminar shear stresses related to specific stacking sequences and in general shear stresses between bonded composite subcomponents and/or hybrid metal/composite subcomponents. A numerical study, performed within the ABAQUS FE code environment, is presented here with the aim to provide a preliminary investigation on the effects of the thermal load on the elastic behaviour of the jointed composite structures representative of a civil aircraft fuselage component. The influence in terms of buckling load and mode variation has been investigated by comparing the numerical results with experimental data from ad ad-hoc thermo-structural test campaign.

In recent years, metallic lattice structures produced by means of the Selective Laser Melting (SLM) process have attracted a growing interest in the scientific community for their valuable stiffness/weight ratio and, consequently, wide range of possible applications. Indeed, additive manufacturing techniques are inherently capable to produce structural configuration with any geometric complexity fulfilling any topological optimization requirement, if compared to the common subtractive production techniques. There are many application fields for lattice structures, such as recoverable lattices, plate lattices or hierarchical lattice structures for applications ranging from heat exchangers, medical implants or acoustic dampers. In many applications, metal lightweight lattice structures are joined with carbon fibers reinforced polymers in order to improve the overall mechanical performances and decrease the weight of structural components. The research activity presented here is focused on the investigation of the feasibility and the effectiveness of hybrid metal/composite lattice structures shock absorbers manufactured by additive techniques. Firstly, an extensive modelling and analysis activity has been carried out focused on the assessment of different configurations of Unit Cells within the commercial FEM code Abaqus, by adopting a simplified Fem model with beam and shell elements formulation The Unit Cells, has been designed as a solid base model, represented by a metallic lattice Repetitive Volume Element and composite skins. Different configurations of metallic lattice structures with and without composite skins have been investigated in order to evaluate their energy absorbing capability under a 6J impact analysis. Once selected the best configuration in terms of energy absorption capability, a refined numerical model with 3D elements formulation and by discretising the composite at ply level has been defined. The mechanical behaviour under impact and the energy absorbing capability have been assessed in order to evaluate the applicability of this hybrid composite panels as multipurpose shock absorber.

Carbon fibre reinforced polymers (CFRPs) are widely employed in aeronautical structural load-bearing components, thanks to their properties in terms of strength, stiffness, weight, and corrosion resistance, if compared to the commonly used metallic materials. Among the other failure mechanisms, fatigue is one of the main causes for CFRPs components collapse. Indeed, when subject to cyclic loading, CFRPs exhibit gradual degradation of the mechanical properties as a result of microcracks propagation. In this work, the empirical model, proposed by Shokrieh and Lessard, for the evaluation of the fatigue induced strength and stiffness degradation, has been implemented in the commercial Finite Element Code ANSYS MECHANICAL, by means of Ansys Parametric Design Languages (APDL). A cycle jump strategy, named SMART CYCLE strategy, has been introduced in the numerical model to avoid the simulation of every single cycle and save computational resources. The SMART CYCLE routine is able to predict cycles where fatigue failure criteria are likely to be verified and to limit the numerical simulation to these cycles where damages propagation in terms of fibre and matrix breakage is expected. The proposed numerical approach has been validated on 30° fibre oriented unidirectional coupons subjected to tensile-tensile fatigue loading condition. The numerical results have been compared with experimental data in terms of number of cycles at failure for different percentage of the static strength. Then, the proposed numerical methodology has been used to investigate the fatigue behaviour of a quasi-isotropic open hole composite panel under compression-compression fatigue loading conditions

In this paper, the damage mechanisms of reinforced composite panels subjected to symmetrical and asymmetrical flexural loading conditions have been investigated. The composite components are representative of a regional aircraft fuselage. Three-point bending tests numerical simulations have been used to assess the influence of the different test parameters on the damage behavior of the investigated component. Then, the most representative configuration has been selected for the experimental bending test. the outputs from the numerical simulations, in terms of stiffness and damage onset and propagation, has been employed, in combination with the experimental data, to accurately describe the damage mechanisms associated to the asymmetric application of the load.

This paper presents the damage tolerant design and verification of a composite materials wind turbine blade expected to be manufactured with the manufacturing process named OneShot Blade® technology. This technology allows the production of wind turbine blades without adhesives and/or bonding processes, leading to a significant reduction in labour hours, costs and materials. Here, the OneShot Blade® oriented design of a 10-meter long fibreglass blade is introduced. Two different configurations (conventional and lightened) have been investigated highlighting their damage tolerant characteristics. Structural performances have been evaluated to verify that the structure complies with the IEC 61400-2 and Germanischer-Lloyd (GL) regulations by considering several loading conditions. Finally, comparisons against a similar wind turbine blade, manufactured by means of a standard process, has been presented, to highlight the advantages of the proposed technology.

For the past six years, we have been benefiting from the discovery by Tsai and Melo (2014) that the trace of the plane stress stiffness matrix ( tr ( Q ) ) of an orthotropic composite is a fundamental and powerful scaling property of laminated composite materials. Algebraically, tr ( Q ) turns out to be a measure of the summation of the moduli of the material. It is, therefore, a material property. Additionally, since tr ( Q ) is an invariant of the stiffness tensor Q , independently of the coordinate system, the number of layers, layup sequence and loading condition (in-plane or flexural) in a laminate, if the material system remains the same, tr ( Q ) = tr ( A ∗ ) = tr ( D ∗ ) is still the same. Therefore, tr ( Q ) is the total stiffness that one can work with making it one of the most powerful and fundamental concepts discovered in the theory of composites recently. By reducing the number of variables, this concept shall simplify the design, analysis and optimization of composite laminates, thus enabling lighter, stronger and better parts. The reduced number of variables shall result in reducing the number and type of tests required for characterization of composite laminates, thus reducing bureaucratic certification burden. These effects shall enable a new era in the progress of composites in the future. For the above-mentioned reasons, it is proposed here to call this fundamental property, tr ( Q ) , as Tsai’s Modulus.

In this work, the crashworthiness of the floor subsection of the cargo area of a civil aircraft is investigated. The structural behaviour under dynamic loading conditions of a structural sub-floor component, made of carbon fibre-reinforced composite material, has been studied in detail from a combined experimental/numerical perspective. Indeed, numerical impact tests characterized by different impact energies have been performed, with advanced numerical tools, finalized to the full understanding of the damage onset and propagation during the impact event. Preliminary numerical quasi-static and dynamic analyses have been carried out on the single sub-components. Then, the impact behaviour of the entire subfloor component has been numerically simulated taking into account the inter-laminar and intra-laminar damages onset and propagation. The comparison between the numerical results and the experimental data from the entire subfloor component has allowed to fully characterize the damage evolution in the investigated component and to provide interesting information finalized to a crashworthy design.

Composite materials structures are particularly susceptible to the damages induced by low-velocity impacts that may result in Barely Visible Impact Damages (BVIDs), which can hardly be identified through visual inspection. These damages are particularly dangerous, since they can critically reduce the mechanical properties of the impacted structures. In this work, the damage induced in impacted long and short fibre composite specimens has been experimentally evaluated by means of Non-Destructive Technique (NDT) inspections. The damages size and location have been evaluated by means of ultrasonic testing to assess the influence of fibres aspect ratio (long and short fibres), fibres material (carbon and glass), volume fraction, and impact energy for low velocity impacts on composite specimens. Considerations about the failure mechanisms arising as a consequence of the impact event and their interactions have been finally introduced.

Composite materials damage behaviour is, nowadays, extensively investigated in the frame of aerospace research programmes. Among the several failure mechanisms which can affect composites, delamination can be considered as the most critical one, especially when combined to compressive loading conditions. In this context, nanofillers can represent an effective way to increase the composites fracture toughness with a consequent reduction of the delamination onset and evolution. Hence, in this paper, the toughening effect of the nanofillers on the delamination growth in composite material panels, subject to compressive load, has been numerically studied. A validated robust numerical procedure for the prediction of the delamination growth in composite materials panel, named SMXB and based on the VCCT-Fail release approach, has been used to perform numerical analyses by considering two different types of nanofillers. Reference material, without nanofillers insertion, has been used as benchmark in order to assess the capability of nanofillers to enhance the fracture toughness in composite laminates.

Interlaminar damage evolution is one of the topics of main interest for composite materials research. An excellent level of knowledge has been achieved in the frame of the studies on delaminations propagation under static loading conditions (as a consequence of low velocity impacts). On the contrary, delaminations evolution under cyclic loading conditions is still an open challenge for the scientific community. Since expensive experimental campaigns are needed to assess the fatigue behaviour of Carbon Fibre Reinforced Polymers (CFRPs), the development and implementation of robust and efficient computational finite elements methodologies has become relevant. In this framework, a robust numerical finite element procedure able to simulate the fatigue driven delamination growth is proposed in this paper. A Paris-law based fatigue tool has been implemented in the commercial Finite Element Code ANSYS MECHANICAL via the Ansys Parametric Design Language (APDL) together with a static delamination procedure (named SMXB) previously developed by the authors. The proposed numerical procedure, based on the Virtual Crack Closure Technique (VCCT), is characterised by load step and element size insensitivity within non-linear incremental analyses. The proposed fatigue procedure, named FT-SMXB, has been preliminary validated at coupon level, by comparing the numerical results to literature experimental data on a unidirectional graphite/epoxy Double Cantilever Beam (DCB) specimen. Actually, the excellent agreement between the achieved numerical results and the literature experimental benchmark data proves the accuracy and the potential of the proposed methodology. Subsequently, a numerical application on a composite stiffened panel with skin stringer debonding under compressive-compressive fatigue loading conditions, has been performed. The Excellent agreement found between the obtained numerical results and available experimental literature data, for this application, in terms of number of cycles to failure and debonding propagation as a function of number of cycles, proves that the use of the proposed numerical tool can be extended, with profit, to geometrically complex and more realistic fatigue test cases.

In the present paper, advanced numerical methodologies have been adopted to investigate the structural behaviour of a composite subcomponent for aerospace applications subjected to quasi-static compression and dynamic loads. The analysed structural component, made of laminated carbon fibres reinforced polymers, is part of the floor support system in the cargo area of a commercial aircraft. The inter-laminar and intra-laminar damage onset and propagation has been preliminary monitored under a quasi-static compressive displacement application. Then, the effects on the structural integrity of two impact energy levels have been analysed: 42 J energy has been applied to study the dynamic behaviour in an elastic linear rate while 585 J energy has been considered to assess the crashworthiness behaviour. The adopted numerical model has been validated by comparisons between the numerical results and analytical mass-spring model results and experimental data in terms of stiffness, strain, and ultimate load. The simultaneous assessment of numerical results and experimental data has allowed to provide a comprehensive insight on the damage onset and propagation leading to the structural collapse of the investigated sub-floor support system.

In this paper, the bonding of carbon fiber reinforced polymer (CFRP) components by induction heating has been numerically simulated. Being the induction heating bonding phenomenon characterised by an intrinsic multi-physical nature, three different numerical models taking into account, respectively, electromagnetic, thermic and structural aspects, have been introduced and applied to a stiffened composite panel to predict the stringer/skin interface characteristics. The electromagnetic model evaluates the energy loss due to the Joule effect while the thermal model is used to calculate the temperature distribution within the adhesive layer between skin and stringer. Then the structural model predicts the ultimate failure load of the stiffened panel taking into account the degree of cure of the adhesive depending on the temperature distribution at interface between skin and stringers by means of an experimentally determined degradation law. The overall developed numerical methodology has been preliminary validated by comparisons with experimental data in terms of adhesive temperature as a function of time on rectangular composite coupons. Finally, a sensitivity analysis has been performed to evaluate the effects of the geometrical parameters of the stiffened composite panel and process parameters on the degree of cure and then on the mechanical properties of the adhesive at interface between skin and stringers.

Composite materials could be very useful when applied to structural rocket engine components since they can allow significant weight savings thanks to their high specific strength and high specific stiffness. In the present work, a carbon fiber reinforced composite has been adopted to replace the typical heavy metallic closeout structure of a regeneratively cooled thrust chamber of a liquid rocket engine. The composite structure has been considered wrapped over the inner liner of the thrust chamber, made of copper alloys, and provides hoop strength for withstanding the fuel/coolant pressure in the cooling channels. The main aim of the paper is to investigate the influence of the geometry and the thermo-mechanical load on the structural response of the analyzed composite closeout. This study is expected to provide a better understanding of the physical phenomena occurring during the service life of the chamber together with an effective identification of the sizing loads that should be considered in the design phase of the closeout structure.

In this work, a numerical model for the simulation of the onset and propagation of fatigue induced damages through a GFRP panel loaded in tension is presented. The model has been validated by numerical/experimental comparisons in terms of stiffness reduction over the number of loading cycles, considering three different applied loads. Finally, the model has been adopted to study the damage behaviour of a panel with a central notch damage.

The purpose of this study is the assessment of the mechanical behaviour of bonded composite repairs. An original multi-linear material model for the elastic-plastic response of adhesives is used in conjunction with an effective FE progressive damage approach for the computation of adhesive stress state within bonded joints in repaired composite structures. The proposed numerical model has been applied to several bonded joints configurations and the obtained numerical results have been compared to experimental and predicted data available in literature providing an interesting insight on bonded composite repairs. Simultaneously, a first demonstration has been given about the capability of the proposed material model to provide detailed and reliable information on the adhesive stress distribution, failure onset, and crack-path propagation.

In this paper, the shear behaviour of a stiffened composite panel with a notch has been numerically simulated by means of a three-dimensional progressive degradation model, able to take into account the intra-laminar gradual degradation of composite laminates. As a relevant added value with respect to state of the art progressive damage models, the proposed methodology, based on energy balance considerations, on three-dimensional failure criteria and on gradual material degradation laws, allows to take into account the real three-dimensional stress distribution, failure onset, and gradual propagation in laminated composites in a Finite Element Analysis without any mesh dependency. A User Defined Material Subroutine (USERMAT) has been used to implement the proposed methodology within the FE commercial code ANSYS©. The resulting numerical tool has been firstly adopted, for preliminary validation purposes, to simulate the tensile behaviour of an Open Hole specimen, focusing on the mesh independency feature of the proposed methodology. Then, the shear behaviour of a stiffened composite panel with a large notch has been simulated taking into account the damage onset and propagation in the large notch area where the three-dimensional stress distribution drives the failure mechanisms. The outstanding agreement, for these applications, between the obtained numerical results and literature experimental data, in terms of damage onset, damage evolution and final failure, confirmed the effectiveness and accuracy of the proposed model and its applicability to complex composite components with pre-existent sources of damage and whatever loading conditions. Moreover, numerical results allowed to deeply analyse the failure mechanisms driving the damage evolution, providing interesting information of the mechanical behaviour of composite laminates with pre-existent damage and on the relevance of the stress model complexity (two-dimensional vs three-dimensional) on the quality of the numerical predictions.

In this paper, a research activity, focused on the investigation of new reinforcements able to improve the toughness of composite materials systems, is introduced. The overall aim is to delay the delamination propagation and, consequently, to increase the carrying load capability of composite structures by exploiting the fiber bridging effects. Indeed, the influence of fiber bridging related Mode I fracture toughness (GIc) values on the onset and propagation of delaminations in stiffened composite panels, under three-point bending loading conditions, have been experimentally and numerically studied. The investigated stiffened panels have been manufactured by using epoxy resin/carbon fibers material systems, characterized by different GIc values, which can be associated with the material fiber bridging sensitivity. Experimental data, in terms of load and delaminated area as a function of the out-of-plane displacements, have been obtained for each tested sample. Non-Destructive Inspection (NDI) has been performed to identify the debonding extension and position. To completely understand the evolution of the delamination and its dependence on the material characteristics, experiments have been numerically simulated using a newly developed robust numerical procedure for the delamination growth simulation, able to take into account the influence of the fracture toughness changes, associated with the materials&rsquo, fiber bridging sensitivity. The combined use of numerical results and experimental data has allowed introducing interesting considerations of the capability of the fiber bridging to substantially slow down the evolution of the debonding between skin and reinforcements in composite stiffened panels.

Composite laminates are characterized by high mechanical in-plane properties and poor out-of-plane characteristics. This issue becomes even more relevant when dealing with impact phenomena occurring in the transverse direction. In aeronautics, Low Velocity Impacts (LVIs) may occur during the service life of the aircraft. LVI may produce damage inside the laminate, which are not easily detectable and can seriously degrade the mechanical properties of the structure. In this paper, a numerical-experimental investigation is carried out, in order to study the mechanical behavior of rectangular laminated specimens subjected to low velocity impacts. The numerical model that best represents the impact phenomenon has been chosen by numerical&ndash, analytical investigations. A user defined material model (VUMAT) has been developed in Abaqus/Explicit environment to simulate the composite intra-laminar damage behavior in solid elements. The analyses results were compared to experimental test data on a laminated specimen, performed according to ASTM D7136 standard, in order to verify the robustness of the adopted numerical model and the influence of modeling parameters on the accuracy of numerical results.

This paper introduces a numerical tool developed for the design and optimization of axial-symmetrical hybrid composite/metal structures. It is assumed that the defined structures are produced by means of two different processes: Additive Layer Manufacturing (ALM) for the metallic parts and Filament Winding (FW) for the composite parts. The defined optimization procedure involves two specific software: ANSYS and ModeFrontier. The former is dedicated to the production of the geometrical and FE models, to the structural analysis, and to the post-process, focusing on the definition of the Unit Cells for the modelling of the metal part. The latter is dedicated to the definition of the best design set and thus to the optimization flow management. The core of the developed numerical procedure is the routine based on the Ansys Parametric Design Language (APDL), which allows an automatic generation of any geometrical model defined by a generic design set. The developed procedure is able to choose the best design, in terms of structural performance, changing the lattice metallic parameters (number of unit cells and their topology) and the composite parameters (number of plies and their orientation). The introduced numerical tool has been used to design several hybrid structures configurations. These configurations have been analysed in terms of mechanical behaviour under specific boundary conditions and compared to similar conventional metal structure.

The presence of cut-outs within thin-walled shell structures is unavoidable, holes being needed for the passage of electrical cables, fuel, or just to reduce the weight of the components. Nevertheless, the high stress concentration can lead to a premature collapse of the structure. For this reason, the preliminary design of cylindrical shell structures with holes needs a profound knowledge of the stress distribution for different loading conditions and constraints. In this paper, a parametric study of a fiber-reinforced composite shell cylinder with an elliptical cut-out has been performed. Three different loading conditions were analyzed: Tension, bending, and torsion. Ansys&reg, script, capable of easily generating and analyzing different geometrical configurations, was used to study the dependence of the geometry on the stress distribution near the cut-out. Finally, graphical and analytical relationships were tentatively extrapolated from numerical results, aimed at linking the geometrical parameters of the cut-out to the maximum stress near the cut-out.

This paper investigates the damage behavior of a reinforced composite panel under symmetrical and asymmetrical bending load. The examined reinforced composite structure is a curved CFRP panel and a co–cured CFRP omega stringer. It may be considered a typical stiffened skin section of a regional aircraft fuselage. The bending loading conditions have been simulated by means of three–point bending tests. The aim of the work is a numerical study, performed to study the influence of asymmetric bending loads on intra–laminar and inter–laminar damage, taking into account the different configurations of the panels and the different orientations of the rollers.

In this paper, the skin&ndash, stringer separation phenomenon that occurs in stiffened composite panels under compression is numerically studied. Since the mode I fracture toughness and, consequently, the skin&ndash, stringer separation can be influenced by the fibre bridging phenomenon at the skin&ndash, stringer interface, in this study, comparisons among three different material systems with different fibre bridging sensitivities have been carried out. Indeed, a reference material system has been compared, in terms of toughness performance, against two materials with different degrees of sensitivity to fibre bridging. A robust numerical procedure for the delamination assessment has been used to mimic the skin&ndash, stringer separation. When analysing the global compressive behaviour of the stiffened panel, intra-laminar damages have been considered in conjunction with skin&ndash, stringer debonding to evaluate the effect of the fibre and matrix breakage on the separation between the skin and the stringer for the three analysed material systems. The latter are characterised by different toughness characteristics and fibre bridging sensitivities, resulting in a different material toughness.

The thermo-structural design of critical components for re-entry vehicles (e.g. nose and wing leading edge) requires the adoption of very detailed and complex finite element thermal and mechanical analyses. Indeed, since very high heat fluxes are expected during the atmospheric re-entry, Ceramic Matrix Composites (CMC) are usually employed because they can guarantee lightness, high damage tolerance, crack/fracture resistance, high tensile, bending and compression strength aside from high temperature stability (1900 °C) and excellent thermal shock resistance. However, such materials, characterized by a complex micro-structure, require demanding numerical tools involving very high computational costs. For this reason, embedded Global/Local and sub-modelling approaches need usually to be employed in order to obtain high level of accuracy, especially in the most critical areas of the components, with a significant saving of the computational cost. These efficient numerical tools allow to perform detailed numerical analyses even in a preliminary design phase where very coarse global models are usually adopted. In the present work, embedded Global/Local finite element models and sub-models are defined with the aim to improve the design of the Wing Leading Edge of an hypersonic vehicle. The aerodynamic heating, experienced along the re-entry trajectory of the CIRA Unmanned Space Vehicle FTB-X has been used as loading condition for the wing leading edge design. Local models have been defined at the interfaces between the internal and the external structure, where structural criticalities are expected, and the corresponding numerical results have been assessed and compared. Finite element commercial environments have been adopted for the numerical computations.

The use of carbon fibers-reinforced plastic materials has progressively increased due to growing requirements of reduction in weight and high resistance. However, for this kind of materials, barely visible damages can become very critical for the carrying load capability during life cycle of primary components. This applies, especially for laminates subjected to impulsive loads such as low velocity impacts. In order to check the damage resistance and tolerance of composite panels subjected to low velocity impacts, compression after impact (CAI) tests are usually performed. In the present work, CAI tests have been carried out on composite plates laminated with plies characterized by two different material systems. According to CAI tests, the specimens have been subject, first, to a low velocity impact with a 20 J impact energy, then a compression test has been performed to assess their residual strength. The experimental results have been discussed and compared.

The work aims to investigate the capability of a vertical tail leading edge to withstand bird strike and to provide general considerations on the improvement of such capability with respect to the mass saving and to the required structural performances by considering different material systems. The assessment of the numerical models was ensured by a double check, and therefore considering the two test cases, aimed to calibrate and validate the numerical models against experimental data, coming from literature: 1) bird strike on metallic flat square plate; 2) bird strike on wing leading-edge with the skin made in metallic sandwich structure (both the honeycomb core and the outer and inner face was in aluminum). Finally, the crashworthy design of several leading-edge configurations, modifying the material and thickness of the skins and the central core of the leading edge, without changing its external geometrical shape, was investigated by adopting the numerical procedures used for the validation test cases. Actually, unidirectional fiber-reinforced composite material systems were considered for the outer and inner face of the skins, while different materials were adopted for the core. A comparison among the developed configurations, in terms of requirement fulfillment, global deformation parameters and weight, was performed.

The design of the wing leading edge of re-entry vehicles is a very challenging task since severe aerothermal loads are encountered during the re-entry trajectory. Hence, advanced materials and structural concepts need to be adopted to withstand the elevated thermal gradients and stresses. Furthermore, particular attention must be paid to the design of hot areas and connections between hot and cold areas of the structure, where the presence of major thermal gradients associated to significant thermal expansion coefficients variations, can lead to damage onset and failure. In order to face this issues, Ceramic Matrix Composites are generally employed as passive hot structures due of their capability to operate at elevated temperatures retaining acceptable mechanical properties. In the present work a novel thermo-structural concept of an hypersonic wing leading edge is introduced and verified by means of an advanced finite element thermo-structural model.

In the last decades, the interest in Shape Memory Alloys (SMA) materials has been considerably increased due to their peculiar morphing capabilities, which can be relevant for many different engineering fields such as automotive and aerospace. Among the characteristics of shape memory alloys, their super-elastic behavior and their capability to recover an initial state, once a deformation has occurred, when subjected to a temperature above a predefined threshold, can be considered the most attractive ones. Possible applications of SMA includes, but are not limited to, switch actuators and morphing structures. In this work, the SMA materials super-elasticity effect and their capability to recover an initial state, due to the heating, are described and numerically reproduced by means of a user material routine implemented in the ABAQUS Standard FEM environment. A test-case representative of a mechanical actuator operated by SMA springs is introduced to assess the robustness of the developed numerical procedure for the simulation of Shape Memory Alloys material behavior.

The development of structures able to autonomously change their characteristics in response to an external simulation is considered a promising research field. Indeed, these structures, called smart structures, can be adopted to improve the aerodynamic performance of air and land vehicles. In this work, an overview and future applications of Shape Memory Alloys (SMA)-based smart structures are presented. The use of SMA materials seems to be very promising in several engineering sectors. Advanced SMA-based devices, designed to improve the aerodynamic performance of vehicles by modifying the shape of the spoiler and the rear upper panel, are briefly introduced and discussed in this paper. Indeed, a simplified model simulating the SMA mechanical behavior has been considered to demonstrate the feasibility of the introduced smart structures for adaptive aerodynamic applications. Numerical simulations of the investigated structures are provided as a justification of the proposed designs.

Crashworthiness is defined as the capability of a structure to guarantee its occupants safety during a crash event. In the present work, the crashworthiness of a composite fuselage section has been preliminary investigated. Indeed, the mechanical behavior of the investigated fuselage section has been assessed by means of the FE code ABAQUS/Explicit, considering an impact event on a rigid surface. The proposed numerical model, capable to take into account the matrix and fiber traction and compression damages, uses a 3D Finite Element formulation to better predict the deformations and the damages onset and propagation during the impact event. Two numerical models have been introduced to study the effect of the passenger floor stiffening on the mechanical response to impact of the entire investigated fuselage barrel. The results in terms of displacements and qualitative deformations of the fuselage, for the analyzed numerical models, have been compared.

In this paper, the debonding behavior of a fiber reinforced composite panel, under flexural loading conditions, is numerically studied. Two material systems, respectively insensitive and sensitive to the fiber bridging phenomenon, are considered. The toughening effect of the fiber bridging phenomenon, on the debonding propagation, is analyzed. An innovative numerical procedure, able to consider GIc variations with the crack length, is proposed. The results of numerical models are compared to evaluate the fiber bridging sensitivity effects on the maximum attained load and on the delaminated area.

The damage detection of composite material structures is one of the major concerns in aerospace field, given their even more frequent use. In this work, the failure of an aerospace omega stiffened CFRP panel subject to compressive load is experimentally and numerically studied. The specimen is characterised by a 0° oriented notch damage, with respect to the load direction, located in the middle of the bay. A compressive mechanical test has been performed to determine the global buckling phenomenon and the progressive fibre-matrix damage development induced by the compressive load acting on the panel. The panel has been instrumented with back-to-back strain gauges in skin and stringers locations. Non-destructive techniques, such as lock-in thermography and ultrasounds, have been used to detect the damage status. Numerical analyses have been carried out by means of the FE software ABAQUS and the results have been compared against the experimental data. Finally, an un-notched configuration has been numerically tested and the results have been compared against the notched configuration, in order to assess the influence of the pre-existent damage on the mechanical and failure behaviour of the structure during service loading conditions.

In this paper, a numerical investigation on the inter-laminar and intra-laminar damage mechanisms, characterising the skin-stringer debonding growth in stiffened composite panels subjected to compressive load, has been carried out. A novel numerical methodology for inter-laminar damage growth simulation, based on VCCT technique and able to overcome mesh and time step sensitivity issues, has been adopted in conjunction with a User Material Subroutine (USERMAT), based on Instantaneous Degradation Models (IDM), to investigate the influence of the intra-laminar damages, in terms of fibers and matrix failures, on skin-stringer debonding evolution. The proposed numerical model has been applied to two configurations of single stiffener composite panels with different geometry and debonding size. The numerical results, in terms of compressive load as a function of applied displacements and debonding evolution up to failure, have been compared to experimental data to assess the effectiveness of the implemented methodology. A further comparison with a previously developed numerical model, taking into account only the inter-laminar damage growth, has allowed assessing the influence of the intra-laminar damage on the skin-stringer debonding propagation and the benefit of the proposed integrated inter-laminar intra-laminar numerical model.

Bird impacts can be extremely critical events for the air transport safety. Since aircraft structures have become more and more complex components, the numerical prediction of the damage onset and evolution induced by a bird impact has become a very challenging task. The aim of this work is to provide a brief overview of the numerical techniques adopted for the prediction of the bird impact phenomenon on a leading edge of a regional aircraft wing. Smooth Particle Hydrodynamics (SPH), Rigid and Lagrangian models have been investigated and the results have been compared and critically assessed.

In the present paper a numerical sensitivity analysis is presented with the aim to assess the effectiveness of the X-FEM method for fracture mechanics applications. Different test cases have been adopted for the numerical analyses to point out the X-FEM method behavior in 2-D and 3-D elastic and elasto-plastic conditions. Comparisons with a standard ductile damage model have been carried out to highlight the advantages of the XFEM method in terms of mesh size and shape independency when simulating cracks propagation.