Preliminary analysis method for FRP laminate impact damage size prediction

Low velocity impact damage on carbon reinforced polymer laminate composites has been identified as a key threat to airframe structural integrity since it reduces the strength under compressive loading. Airworthiness certification specifications dictate that the airframe structural components up to the full scale subassemblies have to adhere to the strength and fatigue airworthiness requirements imposed whilst being damaged. The study presented herein combines a set of numerical tools for generating an approach to numerically quantify the damage size after low velocity impact on FRP laminates

Sensitivity of composite scarf joints to manufacturing deviation and disbond under tensile load

Scarf joints are an effective method of bonding thick composite laminates for applications such as the repair of composite aircraft structures. However, concerns remain about their damage tolerance characteristics. Typically composite scarf repairs to aircraft structures require use of hand tools or rudimentary jigs. If the scarf is incorrectly prepared, this may cause a profile deviation to the joint, affecting the bond line stresses and in turn, reducing the residual strength of the joint or repair. The subject of this work examined the sensitivity of composite scarf joints to machining profile deviation and artificial disbond, when subject to static tensile load. Tensile test specimens were prepared with two different configurations of scarf for representing an undercut or imprecise scarf typical of a machining error. In addition, sensitivity of the scarf joints in the presence of an artificial disbond was also tested. Results indicated that for the specimens tested, the scarf is relatively insensitive to minor profile deviation, but highly sensitive to an artificial disbond. Experimental results were also compared with finite element analysis

Numerical simulation of bolted joints pull through failure

Finite Element Analysis numerical models were generated to simulate and investigate the pull-through damage of bolted joints on composite laminates. Three-dimensional elements were used along with a user material subroutine incorporating the material failure criterion in Abaqus® software. Simplified Discrete Ply Modelling (DPM) and Cohesive Zone Modelling (CZM) were also employed. The numerical model predictive capability was assessed and the parameters influencing the pullthrough failure process were investigated. The residual bearing strength of bolted joints following pull-through damage suggested a qualitative agreement between the numerical model and the experimental testing results

Crashworthiness behaviour of a composite fuselage section with cargo door

This paper present the crashworthiness assessment of a composite fuselage section with a cargo door by means of the Finite Element (FE) software ABAQUS/Explicit. In crashworthiness research, no analysis, either experimental or numerical, of a composite fuselage section with a cargo door has ever been carried out. Therefore, the numerical analysis of the vertical drop test of three models of composite fuselage sections, representative of a regional aircraft fuselage, will be performed: a typical fuselage section without cargo door, a section with the cargo door but not the appropriate reinforced structure and a section with the surrounding door structure reinforced as in commercial aircraft sections with such cut-outs. In order to guarantee the integrity of the passengers as well as the structure, the crash kinematics of each model as well as the accelerations experienced by the passengers have been compared and examined in detail. The comparison between the three models allowed to identify the penalty that a duly reinforced cargo door structure induces on the crashworthiness of a composite fuselage

Experimental testing correlation with numerical meso-scale modelling of CFRP structures and the significance to virtual certification of airframes

The design of structural components has altered fundamentally since laminated composites were proved excellent candidate materials in aerospace applications. The key aspects rendering CFRPs preferable to metals, are mostly their significantly higher specific mechanical properties, and the design flexibility through the stacking sequence selection. However, the currently in use limit and polynomial failure criteria, are inadequate to accurately predict all experimentally observed failure modes and damage specificities of the lamina individual constituents, imposing difficulties in the numerical certification of airframe composites. Thus, component and lamina-level testing are sometimes inevitable, requiring industrial resources which are expensive as well as environmentally costly. For that reason, virtual testing could be more promising in substituting real experimental testing, if conducted under advanced failure criteria which better describe the nature of failure. In this study, the open hole tensile (OHT) test has been simulated under the LaRC05 phenomenological failure criterion, with embedded strain-based progressive damage material behavior. A relatively common composite material in aerospace structures has been selected, IM7 8552 of Hexcel, to compare the numerical strength predictions with its corresponding experimental values. The simulations carried out are based on a standard test method by ASTM international, which address the standardisation of strength tests of polymer matrix composite laminates. The, model was created in ABAQUS/Explicit under the VUMAT user subroutine. The resulted predictions have been found to well – correlate with the testing data, irrespective the specimen stacking sequence

Numerical FEA parametric analysis of CAI behaviour of CFRP stiffened panels

This paper examined the effect of numerical modelling parameters on the accuracy and computational efficiency of Carbon Fibre Reinforced Polymer (CFRP) stiffened panels under Compression After Impact (CAI). Pristine and damaged CFRP stiffened panels were subjected to compression in Abaqus® software using Cohesive Zone Model (CZM) method. Various case studies were examined and the effect of the stiffness parameters of the cohesive elements was critically assessed. Moreover, the required number of cohesive zones to fully capture the damage mechanisms of the impacted and pristine panels under compressive loading was examined. The results showed that a wrong set of parameters can even lead to neglecting the induced damage and can cause severe convergence problems in the numerical model. The importance of the Overall Meshing Factor (OMF) was highlighted and a user-defined subroutine (USDFLD) was applied to capture the decrease in the load bearing capability of an impacted panel prior to the compressive loading, since CZM was found insufficient for this scope. The above-mentioned remarks illustrated the process of investigating the optimum numerical parameters set to achieve an accurate and efficient finite element modelling of the stiffened panels structural performance and maximum load-carrying capability, when subjected to CAI loading

Bird strike virtual testing for preliminary airframe design

Abstract Purpose – The purpose of this paper is to use numerical methods early in the airframe design process and access the structural performance of wing leading edge devices made of different materials and design details, under bird strike events. Design/methodology/approach – Explicit finite element analysis was used to numerically model bird strike events. Findings – Structural performance charts related to materials and general design details were drawn to explore the design space dictated by the current applicable airworthiness requirements. Practical implications – This paper makes use of the current capability in the numerical tools available for structural simulations and exposes the existing limitations in the terms of material modelling, material properties and fracture simulation using continuum damage mechanics. Such results will always be in the need of fine-tuning with experimental testing, yet the tools can shed some light very early in the design process in a relative inexpensive manner, especially for design details down selection like materials to use, structural thicknesses and even design arrangements. Originality/value – Bird strike simulations have been successfully used on aircraft design, mainly at the manufactured articles design validation, testing and certification. This paper presents a hypothetical early design case study of leading edge devices for appropriate material and skin thickness down selectio

Parametric analysis of cohesive zone model method for CFRP stiffened panel CAI behaviour

This paper examined the effect of the stiffness of the cohesive elements on the accuracy and the computational efficiency of Carbon Fibre Reinforced Polymer (CFRP) stiffened panels under ,Compression After Impact (CAI). Abaqus® software was used and the Cohesive Zone Model (CZM) method was applied to capture the damage initiation and propagation of the panels. Various case studies were examined and the effect of the stiffness parameters of the cohesive elements wascritically assessed. Moreover, the required number of cohesive zones to fully capture the damage mechanisms of the impacted and pristine panels under compressive loading was examined. The results showed that a wrong set of parameters can even lead to neglecting the induced damage and can cause severe convergence problems in the numerical mode

Coexistence of gastrointestinal stromal tumor (GIST) and colorectal adenocarcinoma: A case report

<p>Abstract</p> <p>Background</p> <p>Gastrointestinal stromal tumors (GIST) represent the most common mesenchymal tumors of the digestive tract. Over the last ten years the management of GISTs has dramatically altered but their coexistence with other gasrointesinal tumors of different histogenesis presents a special interest. The coexistence of GISTs with other primaries is usually discovered incidentally during GI surgery for carcinomas.</p> <p>Case presentation</p> <p>We present here, a case of a 66-year-old patient with intestinal GIST and a synchronous colorectal adenocarcinoma discovered incidentally during surgical treatment of the recurrent GIST. Immunohistochemical examination revealed the concurrence of histologically proved GIST (strongly positive staining for c-kit, vimentin, SMA, and focal positive in S-100, while CD-34 was negative) and Dukes Stage C, (T3, N3, M0 according the TNM staging classification of colorectal cancer).</p> <p>Conclusion</p> <p>The coexistence of GIST with either synchronous or metachronous colorectal cancer represents a phenomenon with increasing number of relative reports in the literature the last 5 years. In any case of GIST the surgeon should be alert to recognize a possible coexistent tumor with different histological origin and to perform a thorough preoperative and intraoperative control. The correct diagnosis before and at the time of the surgical procedure is the cornerstone that secures the patients' best prognosis.</p

Numerical analysis failure prediction of CFRP stiffened panels in the context of optimal airframe structural performance

This paper presents the effect of stiffener damage on Carbon Fibre Reinforced Composite (CFRP) stiffened panels subjected to compression, for various stiffener design configurations. Nonlinear finite element progressive damage numerical simulations were used for the analysis. The investigation targeted the percentage decrease of the panel compression strength between the pristine (undamaged) and damaged stiffened panel states. The three designed cases sought, were assuming stiffened panels of the same weight but of different stiffener design. The study aimed at displaying that for CFRP stiffened panels used in aircraft structures and designed to carry loads where material strength could be the driver for the maximum compression loading capacity and not the structure’s resistance to buckling, the stiffener geometry and material damage propagation are some of the major parameters for optimal stiffened panel design. In that regards and for cost saving from expensive testing surveys, nonlinear finite element analysis is a valuable tool for preliminary design studies and optimal design down-selection