1,720,970 research outputs found
Use of PFA to determine design methods for composite stiffened panels with discrete damages
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Assessment of Progressive Failure Analysis Capabilities of Commercial FE Codes
No full textPurpose – The main objective of this work is to assess the current capabilities of different commercial finite element (FE) codes in simulating the progressive damage of composite structures under quasi-static loading condition in post-buckling regime.
Design/methodology/approach – Progressive failure analysis (PFA) methodologies, available in the investigated FE codes, were applied to a simple test case extracted from literature consisting in a holed composite plate loaded in compression.
Findings – Results of the simulations are significantly affected by the characteristic parameters needed to feed the degradation models implemented in each code. Such parameters, which often do not have a physical meaning, have to be necessarily set upon fitting activity with an experimental database at coupon level. Concerning the test case, all the codes were found able to capture the buckling load and the failure load with a good accuracy.
Originality/value – This paper would to give an insight into the PFA capabilities of different FE codes, providing the guidelines for setting the degradation model parameters which are of major interest
Numerical Investigation on the Failure Phenomena of Stiffened Composite Panels in Post-Buckling Regime with Discrete Damages
No full textThe increasing use of composite materials for commercial aircrafts has encouraged research activities in structural modelling, post-buckling response, determination and failure mode characterization, of the aircraft structures made of these materials. In order to predict the full potential of aircraft composite structures, is essential to develop a reliable methodology that overcomes the current failure design approach, that is based on the first ply failure criteria, aiming at predicting the effective collapse load of a composite structure taking into account the propagation and/or accumulation of initial local failures.
In the last years, Progressive Failure Analysis (PFA) methodologies, based on the Finite Element Method (FEM), have been developed and employed to evaluate the ultimate load carrying capability of complex composite structures.
The scope of this work is to evaluate the collapse load of stiffened CFRP wing panels in post-buckling regime, by applying PFA methodology. The panel shown in this work is stiffened with T shaped stringers, and it is representative of the upper surface of a typical regional aircraft wing box. By non-linear methodology, with progressive failure option, is possible to estimate the influence that the decrease of the panel stiffness, due to local buckling onset, has on the panel ultimate failure load. A comparison between the undamaged wing panel and the damaged one has been performed in order to evaluate, under compressive load and in post-buckling regime, the combined effect of the reduction of the panel stiffness and of the damage propagation. In detail, a stiffened CFRP panel without any hole and with holes (notched panel), has been analysed in post-buckling regime by applying PFA capability of MSC NASTRAN® code. Different hole locations have also been considered, in order to predict which panel zones, when damaged, can more affect the panel collapse load
Low Weight Design Of Impact Damaged CFRP Stiffened Panels by New Design Criteria And PFA
No full textThe object of this work is to outline new design criteria and analysis approaches for a low weight design of composite stiffened panels. The common industrial approach to satisfy the current certification requirements (EASA AMC 20-29) for composite structures, based on the application of
high conservative knockdown factors to the material strength properties and/or performing extensive test campaigns, can lead to oversized structures and to an increase in costs and timing. Nowadays a new design methodology, based on the incorporation of SHM (structural health monitoring) systems into composite structures, is under consideration aiming at exploiting the full potential of damaged composite materials in favor of a greater weight reduction. By detecting the damages thanks to SHM systems, the structure could be designed with higher design allowables (more reliable detection of BVID) improving the static strength for a reduced damage size detection. Under this topic, two wing box stiffened panels, one critical at strength and another one critical at buckling, have been sized under static compressive loads according to classical design approaches and criteria (reference panels). In the first part of the work a sensitivity analysis, finalized to assess the influence of BVID allowable on the panels’ weight, have been performed. The two reference panels have been re-designed releasing the BVID allowable both on the whole panels and on some of their subcomponents (skin, stringer, etc.), in order to evaluate the weight reduction that could be potentially reached by reliable SHM systems. The results of these analyses provide fundamental requirements for the SHM system definition in terms of “which parameters needs to be monitored and where”. Successively, in order to exploit the effective residual strength of impact damaged panels, progressive failure analysis has been performed considering a discrete damage model against the traditional design approach that is based on the first ply failure design criteria on uniformly damaged panel. PFA has been performed on the panel considering a new simplified design model of BVID by simulating this kind of damage with an equivalent hole. This approach will allow to simplify, in the future, the numerical models simulating the low energy impact effects with high computational time savings. It will be possible to avoid the implementation of induced damage models to determine the laminate residual strength after impact, but simply model the laminate with the appropriate hole in the same BVID location. Some results of this approach are shown in this work supported by some impact and CAI tests on different layups and thicknesses
An Application of SOL 400 to Support the Design of Damaged CFRP Stiffened Panels
No full textThe aim of this work is to show an application of the SOL 400 of MSC Nastran® in order to investigate the final failure
response of damaged composite stiffened panels in post buckling regime, by using progressive failure analysis (PFA)
methodology. This methodology has been applied in order to support the design of composite stiffened panels by
predicting the initiation of the local failure and its propagation up to the final collapse of the panel, in presence of local
damage (barely visible impact damage, BVID) and in post-buckling regime. Discrete damages have been considered in
the skin of the panel. According to the indications enclosed in many guidelines for the preliminary design of composite
structures, a simplified model of BVID has been considered in this work, in particular by simulating this kind of
damage with a hole of 1/4 inches in diameter. The collapse load of the panel has been evaluated both for different
locations of a single damage and for multi-damage scenarios. The results of the analyses illustrate the combined effect
of the reduction of the panel stiffness and of the damage propagation, and also the sensitivity of the residual strength of
the panel with respect to different damage locations and damage density
An Integrated Procedure to Optimize the Design of a Composite Wing
No full textCarbon Fiber Reinforced Plastics (CFRP) composites have demonstrated to be particularly suitable for aerospace structural applications due to their high specific strength and stiffness. The development of innovative methods for the structural optimization of aerospace composite components is a topic of interest for research and industry. In particular newer and fast methodologies are desirable in order to support the preliminary design of aerospace composite structure. This work proposes an integrated multilevel procedure aimed at analyzing and optimizing the design of a composite wing box by means of linear analysis. The target is to make available in a preliminary design phase a reliable structural sizing of a composite (or sandwich) wing by means of a rational and computationally efficient approach without using FEA. Assigned the external loads acting on the wing, its geometry and the design material values, this procedure attempts to achieve the objective of weight reduction by optimizing the rib spacing, the skin/web thicknesses, the layups, the sandwich core height, the spar cap area as well as the stringer one according to a no-strength failure and no buckling onset criteria. The proposed methodology is based on the use of in-house developed codes integrated with commercial codes; among the commercial codes modeFrontier® and Hypersizer® are used. The first one is a multidisciplinary and multi-objective software, while the second one is a design, analysis, and optimization code for composite structures. The main theoretical assumptions of the design methodology foresee the wing conceived as concentrated elements: the skin is sized by bending and shear; the spar webs by shear and the caps by bending. Elementary theories are also applied to determine the internal loads acting on the different structural wing box components, both for stiffened and sandwich panels. The methodology allows an automated design process to be executed with very short computational times offering the possibility to perform sensitivity analyses to study the influence of several factors (wing thicknesses, panel structural concept, stacking sequences) on the structural weight. In this way, the use of iterative FE analyses, which are unsuitable for preliminary design, is avoided. The proposed methodology is a very fast tool to optimize the ply thickness and the final laminate thickness according to minimum weight requirement and minimum gauge constraints that can influence the final material selection and its form (prepreg, fabric). The preliminary wing, sized and optimized with this procedure, represents a reliable sized structure able to withstand the assigned external loads; it has been then modeled and analyzed by FEM analysis in a second stage of the design process: the FEA results confirmed that the final FE structure is very close to the preliminary wing structure. In other words, the proposed preliminary design procedure gave accurate results in a very cost effective way if compared with standard optimizations approaches where only general purpose software are considered. The future work of this research activity is to extend this procedure also to other structural items like a fuselage or a vertical fin
Weight Saving on Composite Stiffened Panels by New Design Philosophy
No full textThe common design approach applied by the civil aircraft manufacturers to satisfy the current
certification requirements of composite structure (EASA AMC 20-29), is based on the application of a very conservative design philosophy that leads to oversized structures. A new design methodology is under evaluation based on the incorporation of SHM (structural health monitoring) systems into composite structures aiming at supporting a new design philosophy exploiting the full potential of composite materials. This work proposes the preliminary design of two aircraft wing box composite stiffened panels: one at the wing root and the other one at the wing tip. These panels have been designed according to the current conservative industrial design approach. Then they have been re-designed releasing some of the conservative design criteria, because they were considered resolved by SHM systems: no BVID knockdown factor, no notch material design allowables (only bonded joints and bonded repair are considered) have been applied and post-buckling regime between limit and ultimate load has been allowed. The new design has shown the greatest weight reduction achievable, the design parameters and panel subparts to which the panel weight is more sensitive. This work aims at providing useful indications on the weight saving expected by applying a new design philosophy based on the information coming from a reliable SHM system and which parameters needs to be monitored and where
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