D. Kelly

A damage zone model for the failure analysis of adhesively bonded joints

The design of structural adhesively bonded joints is complicated by the presence of singularities at the ends of the joint and the lack of suitable failure criteria. Literature reviews indicate that bonded joint failure typically occurs after a damage zone at the end of the joint reaches a critical size. In this paper, a damage zone model based on a critical damage zone size and strain-based failure criteria is proposed to predict the failure load of adhesively bonded joints. The proposed damage zone model correctly predicts the joint failure locus and appears to be relatively insensitive to finite element mesh refinement. Results from experimental testing of various composite and aluminium lap joints have been obtained and compared with numerical analysis. Initial numerical predictions indicate that by using the proposed damage zone model, good correlation with experimental results can be achieved. A modified version of the damage zone model is also proposed which allows the model to be implemented in a practical engineering analysis environment. It is concluded that the damage zone model can be successfully applied across a broad range of joint configurations and loading conditions.

On the design, manufacture and testing of trajectorial fibre steering for carbon fibre composite laminates☆

Trajectorial fibre steering of composite tows within composite laminates has been investigated as a means of producing composite structures with increased strength. Strategies to design the fibre trajectories in laminates have included aligning the fibres with principal stress vectors and a newly developed concept of load paths. In the laminate, the steered plies were inter-mingled with unidirectional or fabric plies. A manufacturing technique for the steered plies has been proposed. Applications include a specimen containing an open hole and a specimen with a pin-loaded hole. Specific strength increases of 62 and 85%, respectively, were achieved.

Strength improvement by fibre steering around a pin loaded hole

A fibre steering technique has been applied around boltholes in carbon fibre reinforced epoxy composite laminates to locally enhance the bearing strength of bolted joints. The procedure can precisely place dry tows of fibre on a prepreg fabric following both the tensile and compressive principal stress trajectories around the hole. The bearing test results indicate that fibre steering improved the peak load of the composite bolted joints approximately in linear proportion to fibre addition by weight. The best result achieved an increase for the peak load by a factor of 2.69. The best improvement of bearing strength was by a factor of 1.36 for a specimen reinforced by 3 k fibre tows in tensile principal stress patterns and 6 k fibre tows in compressive principal stress patterns. The bearing strength improved due to significant increase in peak load and moderate change in thickness.

Multi-objective optimisation of composite aerospace structures

This paper describes a procedure for optimising both cost and weight using the cost parameter, ΔC/ΔW, as a primary design driver. The theory behind the algorithms involving the cost parameter will be stated. A comparison between the Pareto method, which is the standard approach for multi-objective optimisation, and this new innovative approach will be made on a simplified aileron structure. Results from the optimisation applied to a Krueger flap will also be presented.

Failure of transversely stitched RTM lap joints

Abstract An experimental evaluation has been undertaken to investigate the effect of transverse stitching on the strength of composite single-lap joints. Balanced single-lap joints were considered, and the lay-up for the adherends was (0/ ±45/90)s. Specimens were stitched with Kevlar® thread in a zigzag pattern and were manufactured by using the resin-transfer moulding (RTM) technique. Experimental results indicated that stitched joints were over 20% stronger than their unstitched counterparts. The scanning electron microscopy study of the fracture surfaces indicated that the failure is mainly dominated by peel and the stitch threads break inside the adherends and then pull out.

A study of fastener pull-through failure of composite laminates. Part 1: Experimental

Abstract Of the extensive number of investigations examining mechanically fastened composite joints, all but a few are limited to consideration of in-plane modes of failure. An experimental and numerical investigation of fastener pull-through failure of composite joints has therefore been undertaken. The experimental program included an investigation of the influence fastener head geometry, laminate thickness, stacking sequence and material system have upon the pull-through loading response. Circular specimens were transversely loaded to pre-determined displacements of the fastener, sections through the specimen taken and their failure mechanisms investigated with an optical microscope. Pull-through failure was found to be characterised by substantial internal damage similar to that observed for low-velocity impacted composite panels. Failure is not evident from inspection of the laminate surfaces. Damage is manifested in the form of a conically distributed network of matrix cracking and delaminations extending through-the-thickness from the fastener head outer edge, directed away from the fastener hole. The internal/barely visible nature of failure represents a significant departure from that generally considered to distinguish fastener pull-through failure. The means by which to increase resistance to pull-through failure are discussed. This research constitutes work performed as part of the Cooperative Research Centre for Advanced Composite Structures (CRC-ACS) task on highly loaded joints.

A knowledge‐based engineering tool to estimate cost and weight of composite aerospace structures at the conceptual stage of the design process

Purpose – This paper seeks to address issues related to the development of a knowledge‐based engineering system for estimating manufacturing cost and weight of a composite structure at the conceptual stage of a design.Design/methodology/approach – The system has been developed in the CATIA V5 knowledge environment and is applied to structures made of composite materials. At the conceptual stage of the design process, a structure is often represented by simple surfaces. The system adds the details necessary to accurately estimate weight and manufacturing cost using geometry and process‐based techniques. Knowledge captured from an expert was used to construct the knowledge base in the system.Findings – It has been found that the system can provide continuous tracking of the weight and cost as the design evolves. Structural FEA and optimisation using MSC.NASTRAN have been integrated into the design process to enables the designer to conduct “what‐if” analyses to explore different design options involving geo...

An evaluation of failure criteria for matrix induced failure in composite materials

Abstract This paper reports on work being undertaken in the Cooperative Research Centre for Advanced Composite Structures Ltd. (CRC-ACS) to develop improved techniques for predicting the failure of composite materials. The procedures being investigated include a maximum strain criterion for fibre failure. For failure of the resin a new approach, which includes determination of the residual stresses due to manufacturing, is being trialed. This work closely parallels the new criteria proposed by Gosse and Hart-Smith [AIAA/CRC-ACS text on composite materials, submitted for publication] and we have subsequently replaced a simple stress criterion for matrix failure with their proposals based on strain invariants. The new procedures are applied to the failure of laminates in bolted joints with complex steered fibre patterns. Thermal residual stress was included to predict the matrix failure of T-section laminates under loads that open the angle between the flanges and the web. Here a transverse tension stress criterion was used.

Fibre steering for a composite C-beam

Fibre steering of composite tows within lamina is being investigated as a means of producing composite components with increased strength or stiffness when compared to the commonly adopted unidirectional and fabric laminates used in industry. The concept of fibre steering and application to a tensile plate with an open hole and a pin-loaded hole is reviewed. The design and manufacture of fibre steering along trajectories for a cantilevered C-section beam representative of an aerospace control surface spar is then described. The results of experimental testing and finite element analysis of the beam are presented. The results indicate that a significant reduction in deflection is possible over an equivalent flange-stiffened design.

A study of fastener pull-through failure of composite laminates. Part 2: Failure prediction

Abstract The experimental study of fastener pull-through failure in composite laminates reported in Part 1 of this paper found pull-through failure to be characterised by substantial internal damage similar to that observed for low-velocity impacted composite panels. Damage is manifested in the form of a conically distributed network of matrix cracking and delaminations extending through-the-thickness from the fastener head outer edge, directed away from the fastener hole. Analysis is conducted in this paper to identify the mechanisms responsible for failure. Finite element analysis indicated high shear stresses at the fastener head outer edge to be responsible for the matrix cracking in this region. Tensile in-plane stresses are the cause of flexural failures found elsewhere in laminates of reduced bending stiffness. Fastener pull-through failure results from the tensile strength of the resin being exceeded. Matrix cracking was found to be the initial mode of failure with cracks aligning themselves perpendicular to the direction of principal stresses. Interply delamination is a secondary mode of failure and represents a propagation of cracking along the path of least resistance. Delaminations are induced due to excessive interlaminar shear and peel strains in the laminate due to through-thickness deformation and matrix cracking respectively. A numerical procedure for the prediction of failure was developed based upon a progressive damage model and a maximum principal strain criterion. Very good correlation between experimental and predicted pull-through failure loads, failure location and failure sequence were achieved. This research constitutes work performed as part of the Cooperative Research Centre for Advanced Composite Structures (CRC-ACS) task on highly loaded joints.