Papers by Author: A. Apicella

Abstract: A SMART Platform is developed based on sensor readings for Structural Health Monitoring of a stiffened composite panel. The platforms main function is divided into three categories: Passive sensing, Active sensing and Optimal sensor positioning. The platform has self-diagnostic capabilities, i.e. prior to its application the health of the sensors and their connection will be checked to avoid any false alarm. Passive sensing results in impact location and force magnitude detection. Active sensing is performed for damage detection. It results in detecting the damage location and severity. Finally the optimal sensor location can be provided given the number of sensors and probability of detection value. This platform is the first step in applying the developed SHM methodologies to real size structures in service load conditions.

Abstract: A multilevel multiobjective platform for structural sizing reproducing the sequence of actions taken during design and structural sizing in industry is presented in this paper. This platform is integrated at two design levels labeled as Preliminary Design Level and Detailed Design Level. The set of design variables can be divided into a group of variables describing the main conceptual layout that affect the dimensions and architecture of the model and a second group of variables influencing the material and mechanical behavior. This kind of approach can be effective if it is possible to separate the constraints that are strongly dependent on the design variables of different design levels.

Abstract: In this paper the application of the finite element method is presented to modeling piezoelectric sensors and actuators for use in structural health monitoring of composite panels. It this demonstrated that FEM can be used to simulate sensorised composite panels and investigate the damage detection capability of sensors.

Abstract: Fatigue test on a full scale panel with complex loading and geometry has been carried out using a tri-axial test machine specifically designed, built and located in the laboratory of the University of Naples. The aeronautical test panel was designed and manufactured by Alenia. The demonstrator is made up of two skins which are linked by a transversal butt-joint that is parallel to the stringer direction. A fatigue load was applied in the direction normal to the longitudinal joint, while a constant load was applied in the longitudinal joint direction. The demonstrator broke up after about 177000 cycles. Subsequently, a finite element analysis was carried out in order to correlate failure events; due to the biaxial nature of the fatigue loads, Sines criterion was used. The analysis was performed taking into account the different materials of which the panel is composed. The output shows good correlation between experimental data and numerical results, predicting the location on the panel exactly where the failure occurred.

Abstract: A special specimen was created cutting a rectangular notched area from the surrounding of the upper left corner of a wide body aircraft door. Then a fatigue traction load was applied in order to induce an MSD crack initiation and propagation. An innovative DBEM (Dual Boundary Element Method) modelling approach was devised, capable of explicitly modelling the different test article layers with their rivet connections even in a 2d approach. The rivets that are close to the propagating crack are coupled with the corresponding holes by non linear contact conditions, and the accuracy improvements are assessed in comparison with a previous linear analysis, in which traction and displacements continuity conditions on the hole-rivet interface had been imposed. The importance of such influence on the simulation precision need to be assessed due to the strong impact that a non linear analysis produces on computational times. For such a complex problem (three different panels, made of different materials, each one with a variable thickness and connected by numerous rivets), experimental crack propagation data were available for the numerical-experimental comparison. With such non linear approach, a significant improvement on the growth rate correlation is obtained, that justify the increased computational effort.

Abstract: Reducing structural weight is one of the major ways to improve aircraft performance. Lighter and/or stronger materials allow greater range and speed and may also contribute to reducing operational costs. Nowadays composite materials are widely used in “primary” structural components such as fuselage, for which contrasting requirements like lightness and structural strength are required, so particular attention is necessary during its design. In this paper a composite front bulkhead, subjected to ultimate pressure load, was examined. The front bulkhead is made of a composite skin, stiffened with seven vertical stiffeners linked through metallic fittings; the whole system is joined to the fuselage by rivets. A Finite Element model was established: the used elements were four nodes shells, simulating composite layers, and two nodes bar elements, simulating rivets; the structure was clamped and a pressure load was applied to the skin. A linear static stress analysis was performed to calculate strains in particular points in which strain gauges or rosettes are placed: the numerical results, compared with experimental ones, show a good degree of correlation. Stress calculations were performed in order to verify the front and rear bulkhead structural safety.

Abstract: A special specimen was created cutting a rectangular notched area from the surrounding of the upper left corner of a wide body aircraft door. This part of the aircraft skin is made of different layers with variable thickness and material (titanium or aluminum). Then a fatigue traction load was applied and some notches were cut in the different layers in order to speed up the crack initiation and reproduce a realistic crack scenario. Such through cracks were monitored during their propagation along the specimen width, in order to have available for the simulation a realistic initial scenario and experimental propagation data useful for the correlation with the simulated crack path and growth rates. In particular an innovative DBEM modelling approach was devised, using a commercial code (BEASY), capable of explicitly modelling the different test article layers with their rivet connections even in a two-dimensional approach. The results of the simulation show a satisfactory correlation with the experimental crack path and growth rates even for such a complex problem: three different panels (one skin with two doublers), made of different materials, each one with a variable thickness and connected through numerous rivets (whose shear stiffness is taken into account for the simulation).

Abstract: Fatigue test on a full scale panel with complex loading and geometry has been carried out by using a tri-axial test machine located in the laboratory of the University of Naples. The aeronautical test panel was designed and manufactured by Alenia. The demonstrator is made up of two parts which are linked by a transversal joint that is parallel to the stringer direction. A fatigue load was applied in the normal direction to the longitudinal joint, while a constant load was applied in the joint direction. The full scale panel was equipped with strain gauges for deformation state measurements. Preliminary static load tests were performed in order to provide deformation measurements for numerical correlation. The outcomes confirmed that the applied load level is consistent with a linear elastic material behaviour. Three intermediate failures occurred before the final one: the first two under a clip, while in the third case a panel frame failed. Finally after about 177,000 cycles the demonstrator broke down. A non linear finite element analysis was also carried out in order to correlate failure events that occurred during the demonstrator testing.