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In this work, the stringer termination debonding on a tensile-loaded hybrid metallic-composite stiffened panel is numerically investigated. The analyzed structure consists of an omega-reinforced CFRP panel joined with a Z-reinforced aluminum plate. An initial debonding is considered between the composite skin and the omega stringers, and the delamination propagation is investigated as the applied load increases. Subsequently, additional sub-components are introduced to mitigate the delamination propagation. The numerical results are compared and critically discussed.

Composite materials, like metals, are subject to fatigue effects, representing one of the main causes for component collapse in carbon fiber-reinforced polymers. Indeed, when subject to low stress cyclic loading, carbon fiber-reinforced polymers exhibit gradual degradation of the mechanical properties. The numerical simulation of this phenomenon, which can strongly reduce time and costs to market, can be extremely expensive in terms of computational effort since a very high number of static analyses need to be run to take into account the real damage propagation due the fatigue effects. In this paper, a novel cycle jump strategy, named Smart Cycle strategy, is introduced in the numerical model to avoid the simulation of every single cycle and save computational resources. This cycle jump strategy can be seen as an enhancement of the empirical model proposed by Shokrieh and Lessard for the evaluation of the fatigue-induced strength and stiffness degradation. Indeed, the Smart Cycle allows quickly obtaining a preliminary assessment of the fatigue behavior of composite structures. It is based on the hypothesis that the stress redistribution, due to the fatigue-induced gradual degradation of the material properties, can be neglected until sudden fiber and/or matrix damage is verified at element/lamina level. The numerical procedure has been implemented in the commercial finite element code ANSYS MECHANICAL, by means of Ansys Parametric Design Languages (APDL). Briefly, the Smart Cycle routine is able to predict cycles where fatigue failure criteria are likely to be satisfied and to limit the numerical simulation to these cycles where a consistent damage propagation in terms of fiber and matrix breakage is expected. The proposed numerical strategy was preliminarily validated, in the frame of this research study, on 30° fiber-oriented unidirectional coupons subjected to tensile–tensile fatigue loading conditions. The numerical results were compared with literature experimental data in terms of number of cycles at failure for different percentage of the static strength. Lastly, in order to assess its potential in terms of computational time saving on more complex structures and different loading conditions, the proposed numerical approach was used to investigate the fatigue behavior of a cross-ply open-hole composite panel under tension–tension fatigue loading conditions.

Commonly adopted shock absorbers and, in general, crashworthy structural components, based on sandwich structural concepts and/or complex dumping mechanisms, are, generally, characterized by high volumes and significant additional mass. This research activity is focused on the investigation of the feasibility and effectiveness of novel thin additive manufactured hybrid metal/composite lattice structures as lightweight shock absorbing devices for application to structural key components in impact events. These hybrid structures would represent a real step beyond the state of the art of shock absorbers being characterized by an additive manufactured metal lattice core, able to maximize the absorbed energy by plastic deformations and, at the same time, by a composite skin/cohesive coating, fully integrated with the internal metal lattice structure, able to lower the global weight and increase the stiffness and strength of the shock absorber. First, an extensive explicit numerical activity has been performed finalised to the assessment of the mechanical behaviour of basic lattice Unit Cells configurations under impact conditions in shock-absorbing panels. The variation of the geometrical characteristics of the lattice cells have been taken into account by adopting a parametric Python routine in ABAQUS with a simplified FEM formulation based on beam and shell elements. Once identified the key features maximizing the energy absorption capabilities of the metallic core, several complex models with 3d solid element formulation have been developed. A final comparison between the hybrid configurations and the state of the art shock absorbing panels, demonstrated the effectiveness of the proposed lightweight hybrid configuration based on additive manufacturing techniques in terms of mass reduction, mechanical and energy absorption performances .

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%.

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.

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 wing section, subjected to a bird strike event, has been investigated. Indeed, the mechanical behavior of the impacted wing section has been inspected by means of the FE code Abaqus/Explicit. The impact phenomenon has been numerically simulated with a coupled Eulerian–Lagrangian (CEL) approach. The impacted structure has been modeled considering a Lagrangian formulation, while a Eulerian one has been used to model the bird. Moreover, Eulerian–Lagrangian contacts are used for transfer the loads arising from the impact event to the Lagrangian structure. The proposed numerical model, capable to take into account the matrix and fiber traction and compression damages, uses a 3D finite element approach to better predict the deformations and the damages onset and propagation during the impact event. The numerical results, in terms of structural deformation and damage onset and propagation, have been obtained by considering different impact locations and different impact angles.

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 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 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.

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.

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.

Crashworthiness is the ability of a structure to withstand a crash event ensuring the safety of its passengers. The aim of this work is to investigate the damage onset and evolution in a composite fuselage barrel undergoing a vertical drop test on a rigid surface. The mechanical behaviour of the barrel has been assessed by means of a Numerical-Experimental study. Indeed, the experimental data from a full scale barrel test, performed at the CIRA facilities, have been used in conjunction with numerical results, obtained by means of an advanced 3D FEM model, to investigate, in detail, the deformations and the damage development during the crash event. The adopted three-dimensional numerical model, based on an explicit FE formulation, uses Continuum Shell elements and CDM to take into account the onset and the evolution of the intra-laminar damages in each subcomponent of the investigated composite fuselage barrel. The obtained numerical results have been compared with experimental data in terms of accelerations, displacements and deformations to provide a preliminary validation of the adopted FEM model. A special attention has been given to sub-components which demonstrated to mainly influence the global mechanical behaviour of the investigated composite fuselage barrel during the experimental test.

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 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.

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.

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 work, the tensile behavior of a hybrid metallic&ndash, composite stiffened panel is investigated. The analyzed structure consists of an omega-reinforced composite fiber-reinforced plastic (CFRP) panel joined with a Z-reinforced aluminum plate by fasteners. The introduced numerical model, able to simulate geometrical and material non-linearities, has been preliminary validated by means of comparisons with experimental test results, in terms of strain distributions in both composite and metallic sub-components. Subsequently, the inter-laminar damage behavior of the investigated hybrid structure has been studied numerically by assessing the influence of key structural subcomponents on the damage evolution of an artificial initial debonding between the composite skin and stringers.

The aim of this work is the study of the influence of a large notch damage in a stiffened aeronautical panel. In particular, the damage onset and evolution due to a cut-out located in the bay of an omega stiffened composite panel subjected to a compressive load is investigated. Three different cut-outs are considered: parallel, normal, and 45° oriented respect to the load. The effects of such configurations are compared in terms of fibre and matrix failures, in order to better understand which configuration is the most sensitive to these type of damages.

During its service life, an airplane can experience impact events with different foreign objects. In case of impact with a flying bird, the term “Bird Strike” is commonly adopted. Bird strike can be catastrophic, especially when small general aviation airplanes are involved. Indeed, the Federal Aviation Administration (FAA), to prevent catastrophic failures with casualties, obliges airplanes to be able to complete the flight after an impact with a medium dimensions bird. Hence, it is clear that, when designing aerospace components, the potential effects of the bird strike events must be taken into account to guarantee the structural integrity and, as a consequence, the passengers and the pilot safety. In order to optimize the costs, the experimental activity for certification purpose should be suitably prepared and driven by an extensive campaign of numerical simulations. Therefore, the accuracy and the effectiveness of the numerical models able to simulate the bird strike event and its consequences on the structural integrity become of major concern when designing aerospace components. In this paper, the numerical methodologies, commonly adopted for the simulation of the bird strike event, are presented and assessed focusing on their capability to predict the induced damage and the composite components’ residual strength .

The fire behaviour of Polymeric composite structures is one of the most critical aerospace research topics. Indeed, the exposure of Polymeric composite structures to high temperatures leads to material decomposition, associated to thermal and mechanical properties degradation. This degradation causes a reduction of the mechanical performances, which can be of main concern for safety reasons. In this paper, the tensile behaviour of Carbon Fibre Composite Polymer specimens, subjected to fire, has been experimentally and numerically investigated. The material properties degradation has been estimated according to an Arrhenius shape function, which relates the mechanical properties of the composite to the temperature. At first, static structural analyses have been carried out to assess the mechanical behaviour of the investigated specimen without fire effects. Then, a coupled thermo-structural analysis allowed evaluating the fire effect on the specimens’ mechanical and the thermal behaviour. In order to preliminary validate the proposed degradation model, the numerical results, in terms of Load versus Displacements curves, have been compared against data obtained from an ad-hoc experimental campaign where fire condition have been suitably replicated during the mechanical tests.

Natural fibres have been adopted for long time as low-cost fillers in the plastic industry but, nowadays they are starting to replace glass fibres in composite structures thanks to their advantages in terms of weight, cost, and environmental (biodegradability) characteristics. Among the others, one of the most interesting application of natural fibre composites is related to the manufacturing of honeycomb cores for sandwich panels with an improvement in terms of weight saving as well as in mechanical properties and functional capabilities, such as vibration control, heat and energy dissipation. In this work, a numerical model, able to predict the impact behaviour of natural fibres honeycomb cores, has been proposed. An explicit impact analysis has been conducted and the results have been validated by comparisons with experimental data in terms of impact force as a function of impactor displacement. Furthermore, the numerical deformed shapes have been compared with experimental images taken during the impact test. The results have demonstrated the validity and the robustness of the proposed numerical model.

In this paper, a numerical/experimental study is introduced on a carbon fibre reinforced polymer composite panel experiencing the tail-pipe fire phenomenon. This phenomenon, being of main concern for an Airplane during the engine-starting phase, can be qualitatively described as the ignition of a flame outside the nozzle impacting the adjacent aircraft structural components, such as the wing or the tail. The tail-pipe fire phenomenon is characterized by a variable duration and may cause the overheating or even the damage of the aircraft components. A numerical model, able to simulate the thermomechanical behaviour of composite structures under fire, is proposed. The presented approach, considering a strong coupling between the thermal and the structural fields and taking into account thermal and mechanical properties degradation, has been implemented in the commercial FEM software ABAQUS and applied to a stiffened composite panel. The numerical model has been validated by comparing the ABAQUS numerical results to the experimental data obtained by an ad-hoc campaign including mechanical tests under variable thermal conditions. The comparisons, performed in terms of Temperature-Time and Load-Strain evolution histories, showed a good agreement between numerical and experimental data, confirming the robustness of the proposed numerical tool and its effectiveness in predicting damage onset and propagation due to the presence of high thermal gradients.

Carbon fiber reinforced plastic (CFRP) structures are highly sensitive to delaminations, resulting from low energy impacts or manufacturing defects. Non-linear numerical algorithms are mandatory to investigate the complex mechanisms governing the delamination growth phenomena. Although the high computational costs associated to the non-linear algorithms are acceptable in a detail verification design stage, less expensive procedures are desired in a preliminary design stage or during optimization procedure. In this work, a fast numerical procedure, able to determine the delamination growth initiation in composite structures in the framework of a damage tolerant design approach when mixed mode I and II growth is expected, is introduced. The state of the art of the fast delamination growth procedures is critically discussed and improvements to the existing approaches are proposed to extend their applicability and to increase their accuracy. Comparisons with the standard non-linear delamination growth approaches are presented to assess the effectiveness of the proposed novel Fast approach. The results of the proposed fast approach are comparable with the ones obtained by means of standard numerical non-linear technique, allowing up to 95% computational cost saving.

Composite fibre-reinforced materials, under low velocity impacts, can experience simultaneous interacting failure phenomena, such as intra-laminar damage, fibre breakage and matrix cracking, and inter-laminar damage such as delaminations. These failure mechanisms are usually the subject of extensive investigations because they can cause a significant reduction in strength of composites structures leading to premature failure. In the present work, composite plates under low velocity impact are investigated. Experimental data, such as experimental curves and images from non-destructive inspections, are used to characterise the low velocity impacts-induced damage in conjunction with a non-linear explicit Finite element numerical model. The adopted numerical model, implemented in the FE code (Abaqus/Explicit) by a user-defined material subroutine (VUMAT), has been demonstrated to be very effective in predicting the damage onset and evolution and, in general, able to correctly integrate the experimental data by providing useful information about the impact damage localisation and evolution.

In this paper, the compressive behaviour of an omega stiffened composite panel with a large notch damage has been investigated. The influence of intra-laminar and inter-laminar damage onset and evolution on the compressive behaviour of a stiffened panel, characterised by a cut-out located in the middle bay and oriented at 45° with respect to the load direction, has been studied. A numerical model, taking into account delamination and fibre-matrix damage evolution, respectively, by means of cohesive elements and Hashin's failure criteria together with material degradation rules, has been adopted. By comparing the performed numerical analyses, taking into account intra-laminar and inter-laminar damages, the effects of the interaction between delaminations and fibre-matrix damage in the large notch area on the global compressive behaviour of the omega stiffened composite panel have been assessed and critically discussed.

The investigation of the mechanical response of fibre-reinforced composite laminates under impact loads can be very difficult due to the occurrence of simultaneous failure phenomena. Indeed, as a consequence of low velocity impacts, intra-laminar damages, like fibre and matrix cracking, and inter-laminar damages, such as delaminations, can take place simultaneously. These damage mechanisms can lead to significant reductions in strength and stability of the composite structure. In this paper a joint numerical-experimental study is proposed which, by means of non-destructive testing techniques (Ultra-sound and thermography) and non-linear explicit FEM analyses, aims to completely characterise the impact induced damage in composite laminates under low velocity impacts. Indeed the proposed numerical tool has been used to improve the understanding of the experimental data obtained by Non-Destructive Techniques. Applications on samples tested according to the AECMA (European Association of Aerospace Manufacturers) prEn6038 standard at three different impact energies are presented. The interaction between numerical and experimental investigation allowed to obtain an exhaustive insight on the different phases of the impact event considering the inter-laminar damage formation and evolution.

The mechanical response of reinforced composite structures under impact loads is particularly challenging owing to the rising of multiple and simultaneous failure phenomena. Indeed, low velocity impacts may produce intra-laminar damages, like fibre breakage and matrix cracking, and inter-laminar damages, such as delaminations and skin-stringer debonding. As already remarked, these failure phenomena often take place simultaneously, leading to a significant reduction in strength and stability of the composite components. In this paper, the behaviour of stiffened composite panels, with omega shaped stringers, under low velocity impacts is numerically investigated by means of non-linear explicit FEM analyses. Different impact energy levels are considered and correlation with experimental data is provided, in terms of impact force, displacement and energy. A sensitivity analysis has been performed to investigate the influence of numerical models’ approximations on the accuracy of the obtained numerical results. Models with an increasing level of damage simulation details have been adopted to study the effects of combined and separated intra-laminar and inter-laminar failures providing an interesting insight on the modelling requirements for an accurate simulation of the investigated phenomena.

Low velocity impacts induce concurring failure phenomena in unidirectional fiber reinforced composites. Hence a refined methodology able to predict the different failure modes and their interaction is mandatory to correctly predict the damage onset and evolution. Indeed, intra-laminar damage and inter-laminar damage often take place concurrently, causing a significant strength reduction up to composite structure collapse. In this paper, a numerical study is proposed which, by means of non-linear explicit FEM analysis, aims to completely characterize the composite reinforced laminates damage under low velocity impacts by introducing a user defined material model in the FEM code ABAQUS. The proposed 3-D numerical investigation allowed to obtain an exhaustive insight on the different phases of the impact event considering the damage formation and evolution.