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In this paper a structural optimization procedure is applied to an aircraft wing leading edge to improve its capability to withstand bird strike events. First, an experimental-numerical correlation, aimed to calibrate the numerical model against experimental data on a bird strike event, has been performed. In the frame of first step, an estimation of the reliability of the full 3D FEM numerical model adopted to simulate the dynamic event with a reduced error has been given. The onset and the propagation of the damage has been taken into account in order to obtain a realistic deformation of the structure as a consequence of the impact event by means of the Finite Element code Abaqus explicit. Then, the crashworthy design of the leading edge has been performed by adopting the commercial optimization tool ModeFrontier in conjunction with the FE explicit code Abaqus. A dedicated python routine has been used to define a fully parametric simplified wing leading edge model. In order to fulfil the aerodynamic requirement, the external surfaces have been considered fixed during the optimization; therefore, only the internal leading edge components and the skin characteristics has been optimized. As optimization variables, the geometrical properties of the skin and the number and the geometrical properties of the stiffeners have been chosen. The global mass of the model, the stress levels, the maximum deformation and the global stiffness have been considered as constraints. As objective function the energy dissipated due to material failure and to the plastic deformations as been selected. This energy is somehow a measure of the capability to dissipate the impact energy due to a bird strike event. The best configuration provided by the optimization procedure, has been finally compare with the starting configuration in order to assess the added value of the new proposed design in term of energy dissipation and crashworthy structural behaviour.

The use of composite materials aimed to reduce the weight of structural components represents one of the topics of main interest for the aeronautical fields due to the need of designing aircraft components with the minimum possible weight, without compromises on the structural performances. Fuselage includes many stress concentration regions, such as doors and windows. In the present work, the mechanical behaviour of a composite fuselage section is numerically studied, by means of the Abaqus/CAE finite elements code. The load case considered in the frame of simulations is the pressurization load case. The stress distribution near the cut-outs and the crack resulting from the high stress locations have been evaluated and the influence of cu-outs on the pressure and deformation distribution have been assessed.

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 behaviour of the impacted wing section has been investigated by means of the FE code ABAQUS/Explicit. The bird mass has been provided by the certification agencies, and recommended to be limited to 1.82 kg for an impact located on the wings. The impact phenomenon has been numerically simulated by means of a Coupled Eulerian-Lagrangian (CEL) approach. Indeed, the impacted structure has been modelled considering a Lagrangian formulation, while a Eulerian formulation 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 fibre 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. The numerical results, in terms of structural deformation, energy absorption, and damage onset and propagation have been obtained by considering different impact locations and different impact angles

Bird strike represents a source of major risk for aircraft structures. Indeed an impact with a bird during cruise or during take-off and landing phases can lead to serious structural damages threatening passenger safety. Currently, certification tests are required to demonstrate that the airplane can safely operate after a bird strike event. For this reason, the need of accurate numerical models able to predict the structural response under a bird strike event can be of significant help in order to reduce the certification time and costs. In this paper, numerical simulations of a bird impact event on the leading edge of a composite wing section are presented. An approach based on Coupled Eulerian-Lagrangian (CEL) formulation has been adopted and, considering the high velocities of interest, the bird has been modelled as a soft body acting as a fluid on the target structure. The aim of this work is to assess the influence of the adopted numerical model in terms of finite element formulation and failure criteria on the bird-strike phenomenon including impact induced damage onset and evolution. The FE software ABAQUS/Explicit have been used to perform two different numerical simulations with different FE formulations. In the frame of the first simulation, the composite laminates have been simulated by using continuum shell elements and ABAQUS standard composite damage initiation and evolution criteria. In the frame of the second simulation, solid elements (C3D8R) have been adopted in conjunction with a user-defined material subroutine (VUMAT) based on Hashin and Puck damage criteria. The results obtained by means of the two proposed formulations have been compared in terms of damage extension in order to assess the robustness of numerical bird strike simulations and, hence, the dependence of the numerical outputs from the adopted FE formulation.

The adoption of natural fibres as low-cost fillers in the plastic industry has been assessed for long time and, nowadays they are starting to replace glass fibres in composite structures thanks to their advantages in terms of weight, cost, and environmental (biodegradability) behaviour. Among the others, one of the most application of natural fibre composites is in 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 assessed. An explicit impact analysis has been conducted and the results have been validated by comparison 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.

Impact induced damage is ofter of main concern for aeronautic structures. The aircraft can undergo low velocity impacts during the assembly stage and during its operative life. Delamination and other damage mechanisms, such as matrix cracks, fibre-matrix debonding and fiber breakages can be the consequence of impact events with foreign objects under in service conditions and maintenance operations. Dam-ages induced by low velocity impacts can be hardly detected by visual inspections because they often are confined deeply in the laminate thickness. Hence the use of expensive and time consuming non-destructive techniques such as ultrasonic tests, can be required. In order to prevent catastrophic events, as a consequence of impacts, designers usually increase the structural safety margins thereby increasing the weight of the aircraft. In the frame of the present work, in order to quantify the influence of impact events on the load carrying capability of complex composite structures, specific advanced numerical techniques are used to take into account the skin-stringer debonding phenomena arising in stiffened composite panels as a consequence of low velocity impact events.

Local therapy with IL-2 may be very effective in the treatment of different forms of cancer. The aim of this study was to determine the effectiveness of IL-2 locoregional application in the treatment of colon cancer.Twenty eight syngenic BDIX rats were utilized in this study. The rats were divided into two groups of fourteen animals: group T (treatment) and group C (control). All rats of both groups were injected, under the splenic capsule, with T 10(7) DHD/K2/ TRb neoplastic cells. Then, within and around the site of the previous inoculation, the T group was injected with 1 ml of glucosate solutions + 0.1% albumin (BSA) containing 2.5 x 10(6) IU of IL-2 ( Proleukin-Chiron), whereas the C group was injected with 1 ml of BSA alone. After three weeks, rats were sacrificed and the liver and spleen were removed. The following parameters were considered: volume and weight, neoplastic-non neoplastic tissue index of the spleen, mitotic index and vascular density of splenic and hepatic lesions.All the studied parameters showed statistically significant differences in treated and untreated animals.This study of a murine model demonstrated that IL-2 locoregional therapy may be effective in the treatment of colon cancer.

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%.

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.

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.

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.

This work is focused on the investigation of the structural behavior of a composite floor beam, located in the cargo area of a civil aircraft, subjected to cyclical low frequencies compressive loads. In a 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 FE code. Then, explicit non-linear analysis under dynamic loads have been considered, to predict the behavior of the composite structure under cyclic loading conditions.

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.