Applied Composite Materials (v.16, #4)

Entire Life Time Monitoring of Filament Wound Composite Cylinders Using Bragg Grating Sensors: II. Process Monitoring by H. Hernández-Moreno; F. Collombet; B. Douchin; D. Choqueuse; P. Davies; J. L. González Velázquez (197-209).
This article is the second of a series of three papers concerning monitoring of filament wound cylinders using Bragg gratings. In this second part, the tooling presented in Part I is used to embed gratings and thermocouples in filament wound glass reinforced epoxy composite cylinders during fabrication. Bragg grating strain was obtained from wavelength and temperature response, by a calibration technique described here. Results from tests on five cylinders show the Bragg grating’s capability to monitor strain evolution during fabrication, and the capacity to detect several phenomena occurring during cure is established, in addition to quantifying the initial material condition of the cylinder before it enters service.
Keywords: Smart materials; Residual stress; Non-destructive testing; Filament winding

Three-dimensional non-linear Finite Element Analyses (FEA) due to an in-plane loading have been performed to evaluate the out-of-plane normal and shear stresses over the overlap region of a Single Lap Joint (SLJ) on different surfaces. These surfaces have been considered as; (i) two interfacial surfaces between the adherends and the adhesive layer, (ii) the mid-surface of the adhesive layer and (iii) two surfaces beneath the surface ply of both the adherends adjacent to the adhesive layer. The critical locations of onset of adhesion, cohesion and delamination failures on the above mentioned surfaces of the SLJ have been determined using suitable damage criteria. A comparative study due to adhesion, cohesion and delamination failures in the SLJ with Fiber Reinforced Polymeric (FRP) composite adherends have been presented. The effects of simultaneous variations of the delamination positions on the out-of-plane peel and shear stress components have been studied by pre-embedding the delamination damages at the critical locations in both the adherends. It has been observed that the possibilities of onset of cohesion failures in the adhesive layer are higher compared to the adhesion and delamination failures. The detailed analyses showed that secondary peaks of out-of-plane stress components (σ z , τ yz and τ xz ) on the mid surface of the adhesive layer appeared at the locations closer to the delamination fronts due to pre-embedded delamination damages. The highest stress magnitudes on the overlap edge of the SLJ have been reduced significantly when the centers of the delamination damages are exactly aligned with the overlap ends of the joint. No significant variations of stress magnitudes have been noticed either when the delaminations are pre-embedded outside the overlap regions or when the delamination damages are completely entrapped within the overlap region.
Keywords: Adhesion failure; Cohesion failure; Delamination damage; FEA; FRP; SLJ

Test and Modelling of Impact on Pre-Loaded Composite Panels by A. K. Pickett; M. R. C. Fouinneteau; P. Middendorf (225-244).
Currently test and simulation of low and high speed impact of Aerospace composite structures is undertaken in an unloaded state. In reality this may not be the case and significant internal stresses could be present during an impact event such as bird strike during landing, or takeoff. In order to investigate the effects of internal loading on damage and failure of composite materials a series of experimental and simulation studies have been undertaken on three composite types having different fibres, resins and lay-ups. For each composite type panels have been manufactured and transversely impacted under the condition of ‘unloading’ or ‘pre-loading’. For preloading a rig has been constructed that can impose a constant in plane strain of up to 0.25% prior to impact. Results have clearly shown that preloading does lower the composite impact tolerance and change the observed failure modes. Simulation of experiments have also been conducted and have provided an encouraging agreement with test results in terms of both impact force time histories and prediction of the observed failure mechanisms.
Keywords: A- carbon fibres; B- impact behaviour; C- damage mechanics; D- finite element analysis; E-failure criterion

The effect of the helical wood fiber structure on in-plane composite properties has been analyzed. The used analytical concentric cylinder model is valid for an arbitrary number of phases with monoclinic material properties in a global coordinate system. The wood fiber was modeled as a three concentric cylinder assembly with lumen in the middle followed by the S3, S2 and S1 layers. Due to its helical structure the fiber tends to rotate upon loading in axial direction. In most studies on the mechanical behavior of wood fiber composites this extension-twist coupling is overlooked since it is assumed that the fiber will be restricted from rotation within the composite. Therefore, two extreme cases, first modeling fiber then modeling composite were examined: (i) free rotation and (ii) no rotation of the cylinder assembly. It was found that longitudinal fiber modulus depending on the microfibril angle in S2 layer is very sensitive with respect to restrictions for fiber rotation. In-plane Poisson’s ratio was also shown to be greatly influenced. The results were compared to a model representing the fiber by its cell wall and using classical laminate theory to model the fiber. It was found that longitudinal fiber modulus correlates quite well with results obtained with the concentric cylinder model, whereas Poisson’s ratio gave unsatisfactory matching. Finally using typical thermoset resin properties the longitudinal modulus and Poisson’s ratio of an aligned softwood fiber composite with varying fiber content were calculated for various microfibril angles in the S2 layer.
Keywords: Wood fiber composite; Ultrastructure; Microfibril angle; Helical; Cell wall