A Paris Law Based Mesh Independent Numerical Methodology for the Simulation of Fatigue Driven Delamination in Composites

Delamination evolution under cyclic loading is one of the most important research topics for the application of composite materials to aerospace, naval, automotive and, in general, transportation fields. Large experimental campaigns are needed to assess the fatigue behavior of Carbon Fiber Reinforced Polymers (CFRPs), which may result extremely time and cost consuming. Nevertheless, composite materials design needs to take into account the evolution of fatigue driven damage. Subsequently, the development of efficient and robust computational finite element methodologies to evaluate progression of delamination in composite structural components subjected to cyclic loading conditions has become relevant. In this paper, a numerical finite element procedure able to simulate the fatigue driven delamination growth is introduced. A Paris-law based cycle jump strategy, combined with the Virtual Crack Closure Technique (VCCT) approach, has been implemented in the commercial Finite Element Code ANSYS MECHANICAL via the Ansys Parametric Design Language (APDL). The main advantages of the proposed numerical procedure, named FT-SMXB, are related to its independence on the time step and element size in the frame of incremental analyses. The procedure has been preliminary validated, in this research study, at coupon level, by comparing the numerical results to literature experimental data on a unidirectional graphite/epoxy Double Cantilever Beam (DCB) specimen. The significant agreement between the obtained numerical results and the literature experimental benchmark data confirms the accuracy and the potential of the proposed methodology

Investigating the Thermo-Mechanical Behavior of a Ceramic Matrix Composite Wing Leading Edge by Sub-Modeling Based Numerical Analyses

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 °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 °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 °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

On the use of lock-in thermography to monitor delamination growth in composite panels under compression

Abstract The success of composites in automotive, aerospace, and naval applications is mainly related to their aptitude to be tailored to obtain a final product that perfectly fulfills the design requirements. However, during both manufacturing processes and maintenance, some flaws, like delaminations (which may escape simple visual inspection), may be induced in composite structures. The presence of delaminations is of major concern for the load-carrying capability of carbon fiber-reinforced polymer panels. Indeed, delaminations can strongly affect the structural strength and may grow under in-service loads, leading sometimes to catastrophic failures. The aim of this work is to explore the use of lock-in thermography for the monitoring of delamination propagation in composite structures when subjected to generic multiaxial loading conditions. A stiffened composite panel with an embedded skin delamination subjected to compressive loading was taken as a benchmark to assess experimentally the effectiveness of lock-in thermography for monitoring the delamination propagation in situ during the compressive mechanical test. The delamination size as a function of the applied load, observed by lock-in thermography during the execution of the compressive test, was used to validate the results of preliminary numerical computations

Kinematic Approach for a Global-Local Coupling: Compressive Behaviour of a Delaminated Panel:

A kinematic approach to global/local coupling has been applied to investigate the behaviour of a delaminated stiffened composite panel, by using an in-house finite element based procedure. The delamination growth phenomenon has been simulated by employing fracture elements implemented in the B2000++® code, which are based on the Modified Virtual Crack Closure Technique (MVCCT); this technique is able to compute the energy release rate on the delamination front for each fracture mode. A very fine three-dimensional mesh in the delaminated region has been considered in order to obtain a good prediction of the delamination growth. The rest of the structure has been modelled by means of shell elements. A global/local approach based on point-wise multipoint constraint has been implemented in the in-house-code and used to connect shell meshes to solid ones. A numerical application on a delaminated composite stiffened panel taken from literature has been introduced. Models characterized by different levels of complexity, i.e. without delamination, with delamination, with delamination growth have been considered and compared to understand the effectiveness of the introduced kinematic approach

Application of an Additive Manufactured Hybrid Metal/Composite Shock Absorber Panel to a Military Seat Ejection System

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%

Stiffened panels damage tolerance determination using an optimization procedure based on a linear delamination growth approach

The damage tolerance of delaminated composite panels under compressive load is usually numerically evaluated by means of computationally expensive non-linear approaches. In this study, an alternative numerical linear approach, able to mimic the delamination propagation initiation, is proposed. With the aim to exploit its benefits, in terms of computational costs reduction, the proposed linear methodology has been used in this study in conjunction with an optimization analysis to assess the damage tolerance of stiffened composite panels with an impact induced delamination under compression. Indeed, the optimization was aimed to find the minimum delamination growth initiation load for a delaminated stiffened panel with variable delamination size and position, providing indications on the damage tolerance capability of the stiffened panel with an arbitrary positioned and sized delamination induced (as an example) by a low energy impact

Influence of Failure Criteria and Intralaminar Damage Progression Numerical Models on the Prediction of the Mechanical Behavior of Composite Laminates

This work evaluates the effectiveness of commonly adopted local damage evolution methods and failure criteria in finite element analysis for the simulation of intralaminar damage propagation in composites under static loading conditions. The proposed numerical model is based on a User Defined Material subroutine (USERMAT) implemented in Ansys. This model is used to predict the evolution of damage within each specific lamina of a composite laminate by introducing both sudden and gradual degradation rules. The main purpose of the simulations is to quantitatively assess the influence of the adopted failure criteria in conjunction with degradation laws on the accuracy of the numerical predictions in terms of damage evolution and failure load. The mechanical behavior of an open hole tension specimen and of a notched stiffened composite panel under shear loading conditions have been numerically simulated by Progressive Damage Models (PDM). Different failure criteria have been implemented in the developed Ansys USERMAT, together with sudden and gradual degradation rules based on the Continuum Damage Mechanics (CDM) approach. Numerical results have been validated against experimental data to assess the effects of the different failure criteria and damage evolution law on the global mechanical response and local damage predictions in composite laminates

Damage growth in aerospace composites

This book presents novel methods for the simulation of damage evolution in aerospace composites that will assist in predicting damage onset and growth and thus foster less conservative designs which realize the promised economic benefits of composite materials. The presented integrated numerical/experimental methodologies are capable of taking into account the presence of damage and its evolution in composite structures from the early phases of the design (conceptual design) through to the detailed finite element method analysis and verification phase. The book is based on the GARTEUR Research Project AG-32, which ran from 2007 to 2012, and documents the main results of that project. In addition, the state of the art in European projects on damage evolution in composites is reviewed. While the high specific strength and stiffness of composite materials make them suitable for aerospace structures, their sensitivity to damage means that designing with composites is a challenging task. The new approaches described here will prove invaluable in meeting that challenge

On the robustness of finite element procedures based on Virtual Crack Closure Technique and fail release approach for delamination growth phenomena. Definition and assessment of a novel methodology

Numerical procedures based on the combined use of the Virtual Crack Closure Technique and of a fail release approach have been widely used to simulate delamination growth phenomena of composite material structures. This paper starts explaining why this kind of methodologies might not be robust due to mesh and load step size dependency and introduces a novel approach able to cope with the problems identified. Finally the effectiveness and robustness of the proposed procedure, implemented into a commercial finite element software by means of user subroutines, are assessed by comparing the obtained numerical results for a delamination growth phenomenon against literature experimental data on a stiffened panel with a circular embedded delamination under compressive load