Laminates compression failure mode

Sohi. M.M., Hahn, H.T. and Williams, J.G. (1987). The effect of resin toughness and modulus on compressive failure modes of quasi-isotropic graphite/epoxy laminates. In Toughened Composites. ASTM STP 937 (N.J. Johnston cd.). ASTM. Philadelphia, PA. pp 37 60. 

In general, the notched Charpy (ISO 179) [26] and Izod (ISO 180) [27] tests are not meaningful for composites, and ISO 179 recommends that only unnotched specimens should be tested. The difficulty with these tests is that in the notched condition the majority of specimens tested perpendicular to the plane of the test panel delaminate at the root of the notch. This reduces the specimen to a thinner version of the unnotched specimen, which as described above for the interlaminar shear test (see Section 5.4) is susceptible to compression-initiated failures under complex local loads. Specimens cut in the plane of the laminate or sheet will be less susceptible to delamination at the notch tip, and crack growth will be possible from the notch tip. However, other compression shear failure modes are still possible in some composites, and they will not be loaded in this direction in most applications. 

One of the most important properties which control the damage tolerance under impact loading and the CAI is the failure strain of the matrix resin (see Fig. 8.8). The matrix failure strain influences the critical transverse strain level at which transverse cracks initiate in shear mode under impact loading, and the resistance to further delamination in predominantly opening mode under subsequent compressive loading (Hirschbuehler, 1987 Evans and Masters, 1987 Masters, 1987a, b Recker et al., 1990). The CAI of near quasi-isotropic composite laminates which are reinforced with AS-4 carbon fibers of volume fractions in the range of 65-69% has... 

Susceptibility to interlaminar failure is a major weakness of advanced laminated composite materials. It can occur by in-plane shearing (i.e., sliding) (mode II). and out-of-plane shearing (i.e.. tearing) (mode III) as well as by tensile (mode I) deformation. Mode II loading is of particular interest, as values have been shown to correlate with compression after impact data [142.143]. which is required for such purposes as civil aircraft certification. 

In the general approach, the loads are applied incrementally until first-ply failure occurs. The type of failure, matrix or fiber, determines which properties of the failed plies must change to reflect the damage created. This is subjective and can cover a range of possibilities. The most conservative approach would completely discard affected properties for the failed plies. So for fiber failure, E would be set to zero. For matrix failure, E22 and G12 would be set to zero. Then, the loads would be incremented until another ply fails, and the procedure would be repeated to complete failure of the laminate. Less conservative approaches attempt to only partially discount stiffness values of the failed ply and even differentiate between tension and compression moduli. These methods can be reasonably accurate if they are accompanied by selected tests that help better define adjustment factors for the stiffness properties of failed plies. However, they are limited in applicability and accuracy because they are affected by the first-ply failure criterion used to trigger the failure sequence and because they do not correctly capture damage modes such as delamination and the interaction between them such as matrix cracks causing delaminations in adjacent ply interfaces. 

Tension, compression, and shear are the three fundamental modes in which a composite lamina may fail. As the composite material is made up of multiple laminae (layers and plies) of various orientations, the stresses in the lamina s principal directions vary from lamina to lamina. As the load is increased, so do the various stresses in the laminae, and failure values may be attained in a certain lamina in a certain principal direction without the overall laminate experiencing actual failure in other words, the failure of the composite laminate is a progressive phenomenon. This progressive damage evolution is subcritical for a while, but eventually leads to ultimate failure of the composite laminate. 

Figure 20.26 Variation of longitudinal compressive strength of carbon fiber epoxy resin laminate with temperature showing transition from shear mode to buckling mode failure. Failure depends on the shear modulus of the matrix and shear strength of the fibers and a similar effect is observed with the uptake of water. Source Reprinted with permission from Ewins PD, Potter RT, Phil Trans R Soc London, A294, 507-517, 1980. Copyright 1980, The Royal Society of Chemistry.

In a more fundamental vein, Zhang et al. [157] studied the confined crystallization behavior of PLA/acetylated BC nanocomposites prepared by compression molding. The results indicated that acetylated BC favored the crystallization of PLA at higher temperatures. In a similar mode, Quero et al. [158] investigated the micromechanical properties of laminated BC/PLA nanocomposites by Raman spectroscopy as a mean to understand the fundamental stress-transfer processes in these nanocomposites and as a tool to select appropriate processing and volume fraction of the fibers. Results showed that Young s modulus and stress at failure of PLA films were foimd to increase by 100 and 315%, respectively, for 18% volume fraction of BC and BC membranes cultured for 3 days exhibited enhanced interaction with PLA because of their higher total surface area. 

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