No abstract

An analytical and experimental investigation of the failure of selected compression-loaded [ ± 0/ T 0]6* composite laminates is described. A general nonlinear theory is presented for predicting a laminate's compressive strength and failure mode. The theory includes the effects of out-of-plane ply waviness, in-plane fiber waviness, and fiber scissoring. A simple compressive test technique is used to obtain the experimental data. The analytical and experimental results show good agreement for 0<45 deg and show excellent agreement for 0 > 45 deg. The dominant compression failure modes for the laminates in this study were found to be interlaminar shearing, in-plane matrix shearing, and matrix compression. IntroductionE FFICIENT designs using composite materials require a thorough understanding of the mechanisms that affect laminate response. Laminate compressive strength is an important design parameter, and researchers have studied the mechanics of this phenomenon for many years. Dow and Gruntfest 1 postulated that the compressive failure of unidirectional laminates was the result of the fibers buckling within the matrix. Rosen 2 developed a model for unidirectional laminates that focused on fiber instability and predicted the compressive strength of such laminates. Greszczuk 3 ' 4 studied the effects of the constituents on the compressive strength of fiber-and lamina-reinforced composite materials. These studies focused on unidirectional laminates. Several researchers 5 ' 8 have conducted detailed studies of the fiber kinking failure mechanism for compression-loaded unidirectional laminates. Suarez et al. 9 studied failure due to outer-ply instability for compression-loaded multidirectional laminates. Some of their experimental results agreed with predicted strengths when the initial waviness of the outer plies was included in the analysis. Rotem and Hashin 10 and Kim 11 conducted experimental studies of the failure of angle-ply laminates. These studies showed that shearing failure mechanisms significantly affected the failure of angle-ply laminates. A nonlinear theory was proposed in Ref. 12 for predicting the strength of compression-loaded multidirectional laminates. The initial waviness of all plies was included in the theory, and two shearing mechanisms that initiate failure were analyzed: interlaminar shearing caused by initial waviness of the plies and in-plane matrix shearing. The compressive strength of [ ± 0] 5 -class laminates for 0 < 6 < 75 deg was predicted by the theory in Ref. 12. A general, validated theory that quantifies the dominant mechanisms affecting the compressive strength of multidirectional composite laminates (0 < 0 < 90 deg) would be useful for understanding laminate strength.The current investigation was conducted to study the failure of selected compression-loaded multidirectional composite laminates. A general nonlinear theory is described that quantifies the dominant mechanisms affecting the compressive strength for a wide range of undamaged composite laminates. This theory is an impr...

This report evaluates the sandwich beam in four-point bending as a compression test method for advanced composites. To accomplish this evaluation the test method was used to obtain preliminary compressive design data for the temperature range 117 to 589 K (−250 to 600°F) and was analyzed to determine any influence from the honeycomb core on composite mechanical properties. The [08], [908], [(±45)2]s, and [0/±45/90]s laminate orientations were investigated utilizing HTS1/PMR-15 graphite/polyimide composite material. The experimental and analytical results indicated that the test method can be used to obtain compressive elastic constants for graphite/polyimide composites. Elastic data were obtained for a wide temperature range, although many failures were initiated by either load concentrations or top cover debonds. The analysis indicated that the honeycomb core had negligible effects on composite mechanical properties. The top cover carried essentially all the compressive load in the beam test section, and the stress state in the test section of the top cover was essentially uniform and uniaxial.

No abstract

Blade-stiffened, compression-loaded cover panels were designed, manufactured, analyzed, and tested. All panels were fabricated from IM6/1808I interleafed graphite-epoxy. An orthotropic blade stiffener and an orthotropic skin were selected to satisfy the design requirements for an advanced aircraft configuration. All specimens were impact damaged prior to testing. Experimental results were obtained for three- and five-stiffener panels. Analytical results described interlaminar forces caused by impact and predicted specimen residual strength. The analytical results compared reasonably with the experimental results for residual strength of the specimens.