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Delamination

Interlaminar stress in composite structures usually results from the mismatch of engineering

properties between plies. These stresses are the underlying cause of delamination initiation and

propagation. Delamination is defined as the cracking of the matrix between plies. The

aforementioned stresses are out-of-plane and occur at structural discontinuities, as shown in

Figure 4-8. In cases where the primary loading is in-plane, stress gradients can produce an

out-of-plane load scenario because the local structure may be discontinuous.

Analysis of the delamination problem has

identified the strain energy release rate, G, a s a

key parameter for characterizing failures. This

quantity is independent of lay-up sequence or

delamination source. [4-25] NASA and Army

investigators have shown from finite element

analysis that once a delamination is modeled a few

ply thicknesses from an edge, G reaches a plateau

given by the equation shown in Figure 4-9.

where:

t = laminate thickness

= remote strain

ELAM

= modulus beforedelamination

E* = modulus afterdelamination

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Chapter Four PERFORMANCE

Figure 4-8 Sources of Out-of-Plane Loads from Load Path Discontinuities [ASM, En-gineered Materials Handbook]

Figure 4-9 S t ra i n E ne r g y Re-lease Rate for Delamination Growth[OBrien, Delamination Durability ofComposite Materials for Rotorcraft]

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Linear elastic fracture mechanics identifies three distinct loading modes that correspond to

different crack surface displacements. Figure 4-10 depicts these different modes as follows:

Mode I- Opening or tensile loading, where the crack surfaces move directlyapart;

Mode II - Sliding or in-plane shear, where the crack surfaces slide over eachother in a direction perpendicular to the leading edge of the crack; and

Mode III - Tearing or antiplane shear, where the crack surfaces moverelative to each other and parallel to the leading edge of the crack

(scissoring).

Mode I is the dominant form of loading in cracked metallic structures. With composites, any

combination of modes may be encountered. Analysis of mode contribution to total strain

energy release rate has been done using finite element techniques, but this method is too

cumbersome for checking individual designs. A simplified technique has been developed by

Georgia Tech for NASA/Army whereby Mode II and III strain energy release rates are

calculated by higher order plate theory and then subtracted from the total G to determine Mode

I contribution.

Delamination in tapered laminates is of particular interest because the designer usually has

control over taper angles. Figure 4-11 shows delamination initiating in the region A wherethe first transition from thin to thick laminate occurs. This region is modeled as a flat laminate

with a stiffness discontinuity in the outer belt plies and a continuous stiffness in the inner

core plies. The belt stiffness in the tapered region E2

was obtained from a tensor

transformation of the thin region E1

transformed through the taper angle beta. As seen in the

figure's equation, G will increase as beta increases, because the belt stiffness is a function of

the taper angle. [4-25]

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Delamination Marine Composites

Figure 4-10 Basic Modes of Loading Involving Different Crack Surface Displacements[ASM, Engineered Materials Handbook]

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Lately, there has been much interest in

the aerospace industry in the

development of tough resin systemsthat resist impact damage. The

traditional, high-strength epoxy

systems are typically characterized as

brittle when compared to systems used

in the marine industry. In a recent test

of aerospace matrices, little difference

in delamination durability showed up.

However, the tough matrix composites

did show slower delamination growth.

Figure 4-12 is a schematic of a log-log

plot of delamination growth rate,

,

where:

Gc

= cyclic strainenergy releaserate

Gth

= cyclic threshold

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Chapter Four PERFORMANCE

Figure 4-11 Strain Energy Release Rate Analysis of Delamination in a Tapered Lami-nate [OBrien, Delamination Durability of Composite Materials for Rotorcraft]

Figure 4-12 Comparison of DelaminationGrowth Rates for Composites with Brittle andTough Matrices [OBrien, Delamination Durabil-ity of Composite Materials for Rotorcraft]