Date of Award

2014

Document Type

Thesis

Degree Name

Master of Science (MS)

Department

Mechanical Engineering

First Advisor

Shivakumar, Kunigal Dr.

Abstract

A majority of aerospace structures are subjected to bending and stretching loads that introduce peel and shear stresses between the plies of a composite laminate. These two stress components cause a combination of mode I and II fracture modes in the matrix layer of the composite laminate. The most common failure mode in laminated composites is delamination that affects the structural integrity of composite structures. Damage tolerant designs of structures require two types of materials data: mixed-mode (I-II) delamination fracture toughness that predicts failure and delamination growth rate that predicts the life of the structural component. This research focuses on determining mixed-mode (I-II) fracture toughness under a combination of mode I and mode II stress states and then a fracture criterion for AS4/8552 composite laminate, which is widely used in general aviation. The AS4/8552 prepreg was supplied by Hexcel Corporation and autoclave fabricated into a 20-ply unidirectional laminate with an artificial delamination by a Fluorinated Ethylene Propylene (FEP) film at the mid-plane. Standard split beam specimens were prepared and tested in double cantilever beam (DCB) and end notched flexure modes to determine mode I (GIC) and II (GIIC) fracture toughnesses, respectively. The DCB specimens were also tested in a modified mixed-mode bending apparatus at GII m/GT ratios of 0.18, 0.37, 0.57 and 0.78, where GT is total and GII m is the mode II component of energy release rates. The measured fracture toughness, GC, was found to follow the locus a power law equation. 2.08 ) (        = + − T m II C IC IIC IC G G G G G G The equation was validated for the present and literature experimental data.

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