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Applied Composite Materials: An International Journal for the Science and Application of Composite Materials (v.11, #3)
Finite Element Computation of Woven Ply Laminated Composite Structures up to Rupture by Cyril Bordreuil; Christian Hochard (pp. 127-143).
For unidirectional ply laminates, the great diversity of the damage mechanisms and their patterns of evolution make it extremely difficult to estimate the strength margins. In the case of woven ply laminates, the number of damage mechanisms is fairly small (no transverse rupture occurs and the material has a greater resistance to delamination) and the behaviour of the material is fairly simple to model up to rupture. In this study, a numerical model for woven ply laminated composite structures up to rupture is developed. The implementation is performed in a Euler Backward scheme and the consistent tangent stiffness matrix is calculated. Comparison with some experiments on structures are made and the model predicts these experiments well.
Keywords: woven composite; finite element; rupture; damage
Direct Observation of Fatigue Cracking in the Fuel Plate Using the Scanning Electron Microscope by Xi-Shu Wang; Yong Xu; Xian-Qi Xu (pp. 145-154).
Safety and environmental considerations play the important role in selecting and processing fusion materials. Fatigue impairs the reliability of the components utilized in the fusion reactor. In this paper, we described the fatigue cracking mechanism of the sandwich structure of dispersion U3Si2-Al fuel plates using the in situ scanning electron microscope. Direct observations indicated that the failure originates in the vicinal clad-meat interface under tensile-tensile cyclic and three points bending loading. The fatigue crack occurs in two typical fracture modes — Mode-I and the mixed-mode of I and II. The effect of the process of U3Si2-Al fuel meat on the fatigue behaviors of the sandwich structure is obvious.
Keywords: U3Si2-Al fuel plate; sandwich structure; fatigue failure; online; SEM
Integral Manufacturing of Composite Skin-Stringer Assembly and Their Stability Analyses by Hassan Mahfuz; Prasun Majumdar; Mrinal Saha; Frederick Shamery; Shaik Jeelani (pp. 155-171).
Stiffened composite constructions are increasingly being used in the primary structures of aircraft. One key component in these structures is the assembly between the skin and the stringer. The purpose of the stringers sandwiched between two separate layers of skin is to provide structural integrity to a relatively weak skin-structure. Current practice is to fabricate the skin and the stringer separately, assemble them with adhesively bonded joints, and then co-cure the entire assembly in an autoclave. However, the reliability of the joint manufactured in this fashion is not dependable and hence requires riveting of the skin with the stringer by hundreds of mechanical fasteners. Although the mechanical fastener improves the joint reliability, it certainly increases the weight and reduces the strength of the structure by introducing stress concentration points around the rivet holes. In order to eliminate these disadvantages, an innovative low cost manufacturing technique has been developed. The technique utilizes the vacuum assisted resin transfer molding (VARTM) process to co-inject both the skin and the stringer in one integral step. Furthermore, the skin and the stringer are now part of one continuous fabric preform which by default eliminates any adhesive bonding. Several skin-stringer assemblies with plain weave carbon fabric and SC-15 epoxy resin have been manufactured following this procedure. Stability of the manufactured skin-stringer assembly has been investigated experimentally. The extensive analysis focused on the determination of the critical load corresponding to the instability of the structure, failure load and study of the failure mechanisms. Details of manufacturing procedures and experimental investigations are presented in the paper.
Keywords: stiffened composite structure; VARTM; integral fabrication; skin-stringer; stitching; stability analysis
Interlaminar Fracture Toughness of CF/PEI and GF/PEI Composites at Elevated Temperatures by Ki-Young Kim; Lin Ye; Kim-Meng Phoa (pp. 173-190).
An experimental study has been conducted to assess temperature effects on mode-I and mode-II interlaminar fracture toughness of carbon fibre/polyetherimide (CF/PEI) and glass fibre/polyetherimide (GF/PEI) thermoplastic composites. Mode-I double cantilever beam (DCB) and mode-II end notched flexure (ENF) tests were carried out in a temperature range from 25 to 130°C. For both composite systems, the initiation toughness, G IC, ini and G IIC, ini, of mode-I and mode-II interlaminar fracture decreased with an increase in temperature, while the propagation toughness, G IC, prop and G IIC, prop, displayed a reverse trend. Three main mechanisms were identified to contribute to the interlaminar fracture toughness, namely matrix deformation, fibre/matrix interfacial failure and fibre bridging during the delamination process. At delamination initiation, the weakened fibre/matrix interface at elevated temperatures plays an overriding role with the delamination growth initiating at the fibre/matrix interface, rather than from a blunt crack tip introduced by the insert film, leading to low values of G IC, ini and G IIC, ini. On the other hand, during delamination propagation, enhanced matrix deformation at elevated temperatures and fibre bridging promoted by weakened fibre/matrix interface result in greater G IC, prop values. Meanwhile enhanced matrix toughness and ductility at elevated temperatures also increase the stability of mode-II crack growth.
Keywords: interlaminar fracture toughness; temperature effect; carbon and glass fibre; polyetherimide