Kunigal N. Shivakumar

Interlaminar tension strength of graphite/epoxy composite laminates

An L-shaped curved beam specimen and a tension loading fixture were proposed to measure the interlaminar tension strength of laminated and textile composites. The specimen size was 2 X 2 in. (51 x 51 mm). The use of a standard tension test machine and the introduction of load nearly at the specimen midthickness were the advantages of the proposed specimen. Modified Lekhnitskii and beam theory equations for calculating interlaminar stresses of an L-beam were verified by finite element analysis. The beam theory equation is simple and accurate for mean radius to thickness ratios greater than 1.5. The modified Lekhnitskii equations can be used for detailed stress field calculation. AS4/3501-6 graphite/epoxy unidirectional specimens with thicknesses of 16, 24, and 32 plies were fabricated and tested. The delamination initiation site agreed with the calculated maximum interlaminar tension stress location for all three thicknesses. Average interlaminar tension strengths of 16-, 24-, and 32-ply laminates were 47.6, 40.9, and 23.4 MPa, respectively. Results of 16- and 24-ply specimens agreed reasonably well with data in literature for a longer size specimen and a different loading fixture. Interlaminar tension strength decreased with increased specimen thickness and width because of volumetric effect.

Through-the-thickness tension strength of 3-D braided composites

Low cost of manufacturing textile composites make them a candidate for manufacturing affordable aerospace vehicles. However the material test methodologies and database are needed before they can be confidently applied. The study thus far has been on characterization of mechanical properties of two-dimensional (2-D) textile composites. This paper presents through-the-thickness strength test methodology, analysis, and data for 3-D, triaxially braided composites. The braid preform [012k /±β12k ] 45%axial was manufactured using a four-step circular braiding machine and BASF G30-500, 12k graphite fiber tows. The preform was consolidated with Tactix 123 matrix. L-beam specimens were tested in tension and through-the-thickness strength was calculated from the measured failure load and the curved beam equation. Strength of braided composites was compared with an equal thickness and ply stacking AS4/3501-6 graphite/epoxy laminated specimen. The average strength of braided, equivalent laminate, and unidirectional composites was 24.2, 28.8, and 40.9 MN, respectively. Through-the-thickness strength of both braided and laminated composites was found to be a matrix dependent property. Failure modes in braided composites were matrix cracks, tow cracks, and tow separation; whereas failures in laminates were delaminations and ply-cracks. Reduction of size of resin pockets in braided composites may improve the through-the-thickness strength.

Fracture Mechanics Analyses for Interface Crack Problems - A Review

Recent developments in fracture mechanics analyses of the interfacial crack problem are reviewed. The intent of the review is to renew the awareness of the oscillatory singularity at the crack tip of a bimaterial interface and the problems that occur when calculating mode mixity using numerical methods such as the finite element method in conjunction with the virtual crack closure technique. Established approaches to overcome the nonconvergence issue of the individual mode strain energy release rates are reviewed. In the recent literature many attempts to overcome the nonconvergence issue have been developed. Among the many approaches found only a few methods hold the promise of providing practical solutions. These are the resin interlayer method, the method that chooses the crack tip element size greater than the oscillation zone, the crack tip element method that is based on plate theory and the crack surface displacement extrapolation method. Each of the methods is validated on a very limited set of simple interface crack problems. However, their utility for a wide range of interfacial crack problems is yet to be established.

Delamination Fracture Toughness of Woven-Fabric Composites Under Mixed-Mode Loading

25 5 115.63 44.99 i5.164804107807360EC04 11 i5.164804107807630EC04 5 25 6 115.63 44.99 i4.813254329786170EC04 11 i4.813254329786540EC04 5 50 8 263.43 99.15 i4.724964225687430EC05 12 i4.724964225681190EC05 4 50 10 263.43 99.15 i4.319308443921390EC05 12 i4.319308443921450EC05 5 50 12 263.43 99.15 i4.025749894361610EC05 12 i4.025749894361660EC05 5 100 15 596.40 220.70 i3.947662073458550EC06 14 i3.947662073455790EC06 4 100 35 596.40 220.70 i3.016404837142820EC06 15 i3.016404837142080EC06 5 150 50 960.70 353.40 i1.083410906665180EC07 16 i1.083410906665080EC07 5 150 98 961.00 353.00 i1.022353967442190EC07 17 i1.022353967441990EC07 6 150 173 961.00 353.00 i1.119892218088950EC07 20 i1.119892218088840EC07 8 200 196 1347.00 494.00 i2.719539763809190EC07 19 i2.719539763808710EC07 7 200 391 1347.00 494.00 i3.953225968473330EC07 32 i3.953225968463580EC07 12 250 307 1751.00 641.00 i5.987506882057660EC07 21 i5.987506882057300EC07 8 250 571 1750.00 640.00 i9.909334646511780EC07 44 i9.909334646500030EC07 17

EFFECT OF MANUFACTURING PROCESS ON MIXED-MODE FRACTURE TOUGHNESS OF WOVEN FABRIC COMPOSITES

Effect of manufacturing processes on delamination fracture toughness and resistance to growth of T300/934 plane weave fabric carbon/epoxy laminate was evaluated. Three manufacturing processes, namely, autoclave molding (ATM), compression molding (CM), and vacuum assisted compression molding (CMV) were used. Split edge delaminated specimens and mixed-mode bending test apparatus were used to measure delamination fracture toughness. All panels were fabricated and specimens prepared and tested under dry condition and displacement control. Mixed-mode tests for G/G,, ratios of 1/4, 1/1, 2/1, and 4/1 were conducted. Fracture tests were stable for mode-l and G,/GM > 1.0 and unstable for mode-ll and G/GM < 1/4. Average G,. for ATM composites for G/G,, loading of 1/0, 4/1, 2/1, 1/1,1/4 and 0/1 was 282.0, 336.3, 374.8, 390.5, 632.2, and 663.8-J/m2 respectively. Average G0 for CM composites for G/G,, loading of 1/0, 4/1, 2/1, 1/1, 1/4 and 0/1 was 375.5, 381.9, 439.6, 504.4, 746.1, and 933.8-J/m2 respectively. Delamination fracture toughness and resistance to delamination growth of ATM and CMV composites were nearly same for all loading conditions. However, the delamination fracture toughness of CM composite is about 14 to 40 % larger than ATM composites, due to larger resin thickness ahead of the delamination front. In both ATM and CMV, vacuum was used to remove entrapped air and any volatile gases between and within the plies, which resulted in laminated composites having nearly identical flexural modulus and delamination fracture properties.

Some Observations on the Current Status of Performing Finite Element Analyses

Aerospace structures are complex, high-performance structures. Advances in reliable and efficient computing and modeling tools are enabling analysts to consider complex configurations, build complex finite element models, and perform analysis rapidly. Many of the early-career engineers of today are very proficient in the usage of modern computers, computing engines, complex software systems, and visualization tools. These young engineers are becoming increasingly efficient in building complex 3D models of complicated aerospace components. However, the current trends demonstrate blind acceptance of the results of the finite element analyses. This paper is aimed at raising an awareness of this situation. Examples of the common encounters are presented. To overcome the current trends, some guidelines and suggestions for analysts, senior engineers, and educators are offered.