Rodney S. Thomson

The analysis of skin-to-stiffener debonding in composite aerospace structures

Abstract A comprehensive finite element (FE) analytical tool to predict the effect of defects and damage in composite structures was developed for rapid and accurate damage assessment. The structures under consideration were curved, T-stiffened, multi-rib, composite panels representative of those widely used in aerospace primary structures. The damage assessment focussed on skin-to-stiffener debonding, a common defect that can critically reduce the performance of composite structures with integral or secondary bonded stiffeners. The analytical tool was validated using experimental data obtained from the structural test of a large stiffened panel that contained an artificial skin-to-stiffener debond. Excellent agreement between FE analysis and test results was obtained. The onset of crack growth predictions also compared well with the test observation. Since the general damage tolerance philosophy in composite structures follows the “no-growth” principle, the critical parameters were established based the onset of crack growth determined using fracture mechanics calculations. Parametric studies were conducted using the analytical tool in order to understand the structural behaviour in the postbuckling range and to determine the critical parameters. Parameters considered included debond size, debond location, debond type, multiple debonds and laminate lay-up.

Experimental Investigation of Dynamically Loaded Bolted Joints in Carbon Fibre Composite Structures

This paper presents quasi-static and dynamic modelling of bolted composite structures using the explicit finite element code PAM-CRASH. User controlled point link (PLINK) elements were investigated for modelling the bolted composite joints used in the structures. Simulation results were compared with quasi-static and dynamic structural testing reported previously. Two loading configurations were considered. It was shown that the PLINK element modelling approach agreed well with the experimental results for both loading configurations and for one case offered significant improvements over other simplified bolt modelling methods. A stacked shell modelling approach was used to model the interlaminar delamination damage present in the ball-loaded impact mode. The overall response of the structure was significantly improved by the addition of these energy absorbing interfaces.

A Finite Element Methodology for Analysing Degradation and Collapse in Postbuckling Composite Aerospace Structures

A methodology for analysing the degradation and collapse in postbuckling composite structures is proposed. One aspect of the methodology predicts the initiation of interlaminar damage using a strength criterion applied with a global-local analysis technique. A separate approach represents the growth of a pre-existing interlaminar damage region with user-defined multi-point constraints that are controlled based on the Virtual Crack Closure Technique. Another aspect of the approach is a degradation model for in-plane ply damage mechanisms of fiber fracture, matrix cracking, and fiber-matrix shear. The complete analysis methodology was compared to experimental results for two fuselage-representative composite panels tested to collapse. For both panels, the behavior and structural collapse were accurately captured, and the analysis methodology provided detailed information on the development and interaction of the various damage mechanisms.

A Review of Explicit Finite Element Software for Composite Impact Analysis

As explicit finite element (FE) codes improve and advanced material models become available, such tools will find more widespread application within the aerospace industry, as ‘what-if ’ simulations become more manageable with increasing computing power and greater modeling realism. This paper describes the investigation of three commercial explicit FE analysis packages, LS-Dyna, MSC.Dytran, and Pam-Shock, to determine their capabilities in predicting barely visible impact damage (BVID) in composite structures. The investigation is conducted by first determining the suitability of the codes in constructing an FE model of a stiffened panel, solving for BVID and retrieving results. The results are in turn compared to experimental data in order to gauge the suitability of the codes for composite design and analysis. Comparisons of the FE simulations to experimental data include damage development and degradation, as well as the time-history responses. The Chang-Chang failure theory with brittle degradation was used for both LS-Dyna and MSC.Dytran, while the biphase model was used for Pam-Shock. Results indicated that the general shape of the force-time curves as well as the peak forces were predicted reasonably well. However, all simulations predicted a trough that was much less significant than the test results, as well as a shorter impact duration.

A Stacked-Shell Finite Element Approach for Modelling a Dynamically Loaded Composite Bolted Joint Under in-Plane Bearing Loads

This paper presents the results of a study into a novel application of the “stacked-shell” laminate modelling approach to dynamically loaded bolted composite joints using the explicit finite element code PAM-CRASH. The stacked-shell approach provides medium-high fidelity resolution of the key joint failure modes, but is computationally much more efficient than full 3D modelling. For this work, a countersunk bolt in a composite laminate under in-plane bearing loading was considered. The models were able to predict the onset of damage, failure modes and the ultimate load of the joint. It was determined that improved debris models are required in order to accurately capture the progressive bearing damage after the onset of joint failure.

AN ANALYSIS TOOL FOR DESIGN AND CERTIFICATION OF POSTBUCKLING COMPOSITE AEROSPACE STRUCTURES

In aerospace, carbon-fiber-reinforced polymer (CFRP) composites and postbuckling skin-stiffened structures are key technologies that have been used to improve structural efficiency. However, the application of composite postbuckling structures in aircraft has been limited due to concerns related to both the durability of composite structures and the accuracy of design tools. In this work, a finite element analysis tool for design and certification of aerospace structures is presented, which predicts collapse by taking the critical damage mechanisms into account. The tool incorporates a global–local analysis technique for predicting interlaminar damage initiation, and degradation models to capture the growth of a pre-existing interlaminar damage region, such as a delamination or skin–stiffener debond, and in-plane ply damage mechanisms such as fiber fracture and matrix cracking. The analysis tool has been applied to single- and multistiffener fuselage-representative composite panels, in both intact and predamaged configurations. This has been performed in a design context, in which panel configurations are selected based on their suitability for experimental testing, and in an analysis context for comparison with experimental results as being representative of aircraft certification studies. For all cases, the tool was capable of accurately capturing the key damage mechanisms contributing to final structural collapse, and suitable for the design of next-generation composite aerospace structures.

Computer modelling of impact on curved fibre composite panels

This paper presents results of a finite element (FE) analysis study into low energy impact on curved composite panels. The aim of this study was to determine the accuracy and efficiency of a non-linear explicit FE code, MSC.Dytran, and compare results to published experimental data. The study also looked at impact response as a function of composite panel curvature, composite mesh density, impactor weight, velocity and size, and various suggestions are made for improving the accuracy and efficiency of FE analysis procedures in composite low energy impact studies. The paper presents 265 explicit computer simulation results, which show that non-linear FE analysis does provide accurate, efficient and conservative solutions provided various guidelines are followed.

Progressive damage in single lap countersunk composite joints

This paper presents an experimental and computational investigation of the influences of countersink and bolt torque on the progressive failure of single-lap composite joint. Using the Abaqus® software, delamination damage and ply fracture are modelled using cohesive element approach and continuum damage mechanics method, respectively. The model is first validated against a filled-hole tension test to calibrate the composite damage model. Comparison with the experimental results indicates that the computational model is capable of accurately predicting the joint strength and the damage progression process.

Experimental Validation of the 2D Meshless Random Grid Method

This paper presents the first systematic effort for the experimental validation of the 2D Meshless Random Grid Method (MRGM) for the full field measurement of displacement and strain fields. Although the MRGM has been demonstrating very promising characteristics of accuracy, performance and ease of application based on previously conducted sensitivity analysis supported by virtual data, extensive experimental validation was not available until now. This work comes to fill this gap and presents preliminary validation results against strain gauge data collected from open hole tension experiments of composite specimens. In addition, strain and displacement field verification is performed by comparison studies with finite element analysis results.Copyright

Experience with the finite element modelling of a full-scale test of a composite aircraft control surface

A full-scale, co-cured, carbon fibre composite control surface representative of those found on mid-size, jet transport aircraft has been designed and tested. It was designed as a postbuckling blade-stiffened structure to reduce weight and improve operational performance compared with a honeycomb sandwich panel design traditionally used in such structures. The purpose of the tests was to demonstrate the validity of the design methodology by applying static limit and ultimate loads to the structure. The control surface, manufactured from prepreg tape, was successfully loaded to the ultimate design load without evidence of failure. Buckling initiated at approximately 48% of ultimate load with significant out-of-plane displacements observed. The global behaviour predicted by the finite element (FE) model of the test arrangement was in close agreement with the experimental results. Good agreement was also demonstrated with the local behaviour, as evidenced by the strain and buckling results. The inability of the FE analysis to capture complex snap-through mode change behaviour at 89% of ultimate load was identified as a limitation. The success of the testing program demonstrated the suitability of the design methodology for this type of structure.