ASM Metals HandBook Vol. 8 - Mechanical Testing and Evaluation

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Mechanical Testing of Fiber-Reinforced Composites

Dale Wilson, The Johns Hopkins University, Leif A. Carlsson, Florida Atlantic University

Fatigue Testing

Strain gages are commonly used to measure the strain response of the specimens. However, strain gages only

measure local strains. In general, fatigue damage in unnotched composite laminates may be quite random.

Unless the strain gage is quite large, the measured strain may or may not reflect the random damage. O'Brien

(Ref 83) has discussed the effect of gage-section size on composite stiffness measurements. In most cases,

strain measurement over a 25 mm (1.0 in.) gage section is recommended. This will result in a global

measurement of the effect of damage on laminate moduli. There is a trade-off, however; too large a gage length

will not be sensitive to local damage, and too small a gage length will miss local damage.

In lieu of strain gages, clip-on extensometers can also provide excellent accuracy without the difficulty of

bonding a strain gage in place. Most extensometers only give the strain in the axial direction of the specimen,

although there are some special extensometers that can measure the transverse strain in the laminate as well.

Since most extensometers are at least 13 mm (0.5 in.) in gage length, they will measure the global deformation

of the composite. Extensometers are available for operation at a wide variety of temperatures.

As an example, the cyclic stress-strain response of a boron/aluminum laminate presented in Fig. 57 can be

analyzed to yield several mechanical properties. These properties include the following:

• Initial elastic modulus (tangent modulus)

• Stiffness (secant modulus)

• Permanent plastic deformation (strain)

• Elastic unloading modulus after N number of cycles

• Stiffness after N number of cycles

• Cyclic strain

• Number of cycles to fatigue failure

• Strain to static failure

• Residual strength after a set number of cycles

Fig. 57 Schematic of the cyclic stress-strain behavior of a quasi-isotropic laminate of boron/aluminum as

a function of number of cycles. Source: Ref 84

A detailed definition of each property along with the physical significance is presented next.

Initial Elastic Modulus. The elastic modulus E is the tangent modulus of the initial linear portion of the loading

curve. The initial elastic modulus E

is measured on the first loading cycle as shown in Fig. 57. E

usually

correlates with elastic modulus predicted by lamination theory for an undamaged composite. The initial elastic

modulus is significant in that all later measurements of elastic modulus at N cycles will be compared to this

initial modulus to give an indication of the amount of fatigue damage accumulated by the composite laminate.

Stiffness. The stiffness E

is defined as the secant modulus of the loading cycle as indicated in Fig. 58. In order

to eliminate permanent plastic deformation from the stiffness measurement, the stiffness is defined utilizing the

unloading portion of the cycle (the stress range, ΔS, divided by the unloading strain range, Δε). The initial

stiffness can be compared to measurements of stiffness at N cycles and is therefore significant.

Fig. 58 The cyclic stress-strain response of a boron/aluminum ([0/±45/90/0/±45/ ]

, V

= 0.45) composite

showing a dramatic change due to fatigue damage and cyclic hardening of the matrix material. (a) After

4 cycles. (b) After 500,000 cycles

The stiffness may be a useful measurement to design engineers since this is the total strain of the material in

response to an applied load. (This is after no additional permanent plastic strain occurs during cyclic loading.)

The stiffness is a combination of the elastic properties of the laminate as well as the plastic behavior of the

metal matrix.

The stiffness is, therefore, a function of the matrix yield stress. The matrix may cyclically harden and the

stiffness increase during the time damage initiates. (In some metals/heat treat conditions, the material may even

cyclically soften. The properties of polymer matrix materials can also change with cycling and

time/environmental dependent aging.) However, as the fatigue damage in the laminate grows, the stiffness

decreases. Therefore, one cannot determine the onset and extent of fatigue damage in a metal matrix composite

using a stiffness (secant modulus) criterion alone.

Permanent Plastic Strain. The permanent plastic strain, ε

, is measured directly from the stress-strain curve. The

permanent strain is usually accumulated in the first three or four cycles with little additional permanent strain

accumulating thereafter. The accumulated permanent strain gives a good indication of the amount of residual

stress in the 0° fibers since all the constituents in the composite are assumed to strain the same.

Elastic Unloading Modulus after N Number of Cycles. The elastic unloading modulus E

is measured directly

from the stress-strain curve at intervals throughout the life of the composite (Fig. 58). The percentage of the

elastic unloading modulus at N cycles E

as divided by the initial elastic modulus:

× 100 = %E

can be plotted against the number of fatigue cycles N in order to present fatigue damage as a function of

constant amplitude cycling at a prescribed stress ratio and maximum stress.

Stiffness After N Number of Cycles. The stiffness (secant modulus) is measured as the stress range divided by

the unloading strain range. The percentage of the initial stiffness is calculated by dividing the stiffness at N

cycles E

by the initial stiffness E

× 100 = %E

The % E

can be plotted against N in order to compare a “stiffness” and “elastic modulus” criteria for assessing

fatigue damage.

Cyclic Strain. The cyclic strain Δε is measured directly from the stress-strain curve. This strain range can be

used to examine the possible correlation between the composite strain and fatigue endurance. The cyclic strain

can also be multiplied by the elastic fiber modulus to calculate the cyclic fiber stress in the 0° plies.

Number of Cycles to Fatigue Failure. The number of cycles to fatigue failure N

plotted against the maximum

cyclic stresses for a given value of stress ratio is the traditional S-N curve. This type of data can be used to

determine the fatigue endurance limit stress and the maximum cyclic stress at a given stress ratio at which the

composite laminate can be expected to survive beyond some endurance life. The S-N curve does not, however,

indicate the physical condition of the laminate at the established endurance limit other than the fact that the

laminate can survive the endurance stress.

Residual Static Strength. Fatigue specimens that survive a given number of cycles without failure can be tested

for residual strength to correlate fatigue damage and residual strength. The stress at failure, which is the

maximum stress from the stress-strain curve, is defined as the residual static strength S

Strain to Static Failure. For static strength tests, unfatigued composite specimens are loaded to failure. Thus, the

strain to static failure ε

ftot

is used to estimate the 0° fiber ultimate stress.

Footnote

The section “Interlaminar Shear Properties of Fiber-

Reinforced Composites at High Strain Rates” was written

by John Harding and Stephen Hallett, Oxford University. The section “Fatigue Testing and Behavior of Fiber-

Reinforced Composites” was written by W. Steven Johnson and Ramesh Talreja, Georgia Institute of

Technology.

References cited in this section

83. T.K. O'Brien, Effect of Gage-Section Size on Composite Stiffness Measurements, Compos. Technol.

Rev., Vol 1 (No. 4), 1979, p 5–6

84. W.S. Johnson, Fatigue Testing and Damage Development in Continuous Fiber Reinforced Metal Matrix

Composites, ASTM STP 1032, 1989, p 194–221

Mechanical Testing of Fiber-Reinforced Composites

Dale Wilson, The Johns Hopkins University, Leif A. Carlsson, Florida Atlantic University

Damage Assessment

The common methods of assessing the extent and type of damage can be divided into those techniques that are

nondestructive and those that are destructive. The manual edited by Pendleton and Tuttle (Ref 85) is an

excellent reference for several of the techniques, which are described further in the following.

Nondestructive Techniques

Edge Replica Technique. In the edge replica technique, a permanent impression of the specimen edge is

produced in a cellulose acetate film. The advantage of this technique is that the replica can be taken while the

specimen is under load in the test machine. The replica can then be examined later for details of damage, such

as matrix cracks, split fibers, or fiber/matrix separation. This technique can be easily used for replicating any

surface for microstructural detail.

A softened acetate film is held against the desired surface to be replicated. Acetone is applied to the face of the

film to soften it. The acetone serves to polymerize the acetate, making it viscous enough to flow into the

microstructural cavities present on the surface. Pressure is applied to the backside of the film to aid in the flow

process. The acetate film is allowed to dry in place for 3 to 5 min and then is carefully peeled from the surface.

The replica can now be sputtered with gold (a thin layer of gold-palladium alloy) and viewed in a scanning

electron microscope (SEM).

Radiography. As discussed earlier, much of the damage that develops in composite materials is of an internal

nature (i.e., not readily observable from the surface). One of the more popular techniques to assess this internal

damage is radiography. There has been much written about this technique, and a good summary paper can be

found in Ref 85.

Depending on the differences in the x-ray absorptivity between the fiber and matrix, differing details of damage

can be determined. Most fibers do not show well in radiographs, and it is nearly impossible to see individual

fiber failures in graphite/epoxy composites. However, solutions for penetrant enhancement can allow for clear

identification of matrix cracks and delaminations. The penetrant has a significantly higher opacity than the

composite, so areas into which it penetrates show up well in the radiograph. For example, zinc-iodine solution

enhanced the display of ply cracking and delamination in the radiograph of a composite after static testing (Ref

86). With penetrants, damage must come to a free surface (the specimen edge or radiate from a notch) so that

the penetrant can wick into the damage areas. This is usually never a problem since most damage does start at

the free edge or from a notch.

Several laboratories have special rooms set up such that specimens can be radiographed while under load in the

testing machine (Ref 86). This makes the study of damage progression much more efficient.

Acoustic Emissions. This technique can be used real time to “listen” to damage occurring. Each noise emission

has an energy level associated with it. Through much experience with a given composite system, the energy

level of a given event can be associated with a damage mode (See Ref 87). The number of events or acoustic

emission counts is also used to correlate with the extent of damage. This technique is very indirect, requiring a

lot of experience and interpretation of data. Once again, Ref 86 has a good description of this technique.

C-scans. Ultrasonic through transmission C-scans are very popular for assessing initial quality of composite

materials and components. They are an excellent way to assess the presence of delaminations, air pockets, or

significant changes in density (for example, resin rich areas). The initial quality as indicated by C-scans may

provide a good indicator of how the particular laminate will behave under fatigue loading (Ref 88). This

technique is a little less useful for monitoring damage evolution since the part or specimen must be removed

from the machine/structure and placed into the C-scan tank.

Destructive Techniques

Micrographs. After the specimen fails, the fracture surfaces and interface characteristics can be examined using

SEM and energy dispersive x-ray analysis (EDXA). A SEM can be used to distinguish fatigue striation

markings from dimple fracture patterns in order to determine if a portion of the fracture surface failed because

of fatigue or static fracture. The SEM is also very useful for close-up viewing of any type of microstructural

damage. The EDXA is useful for the determination of the chemical composition of the failure surface. This is

especially useful for determining whether the failure occurred in the fiber, matrix, or in some reaction zone

between the fiber and the matrix. Taking photographs through a 30 to 60 power optical microscope has also

been used with considerable success to document damage, such as matrix cracking.

Matrix Etching. After a MMC specimen has been statically or fatigue loaded, matrix etching can be used to

reveal two types of composite damage. First, etching away the matrix material will expose the fibers. The fibers

can then be inspected for breaks. Second, by carefully etching away the matrix in the outer ply and then

removing the exposed fibers, the next layer can be examined for matrix cracking and fiber damage. This

technique can be used to systematically examine the damage of each ply in a laminate.

For 6061 aluminum matrix composites, a 30% hydrochloric acid (HCI) solution in distilled water was found to

be effective for etching away the matrix without detectable damage to either boron or silicon-carbide fibers.

The actual etching process must be carefully controlled if one wishes to remove one layer at a time to inspect

the matrix in the underlying layers. Enough matrix material must be etched away to allow the surface layer of

fibers to be removed; however, if too much matrix material is removed, the matrix damage in the underlying

ply will be etched away. This process is a trial and error operation at best and often requires several attempts

before a satisfactory result is obtained.

Laminate Deply. The specimen-deply technique is a simple destructive-examination method whereby the

individual plies of a polymer matrix laminate are separated for easy examination (Ref 85). This technique is

excellent for determining the laminate-stacking sequence, studying the details of broken fibers, and determining

the extent and precise interface location of delaminations. The interlaminar-bond strength is broken down by a

partial pyrolysis of the matrix resin between the plies of the laminate. This allows for the separation

(unstacking) of the individual plies, provided the adjacent plies have different fiber angles. The individual plies

may then be microscopically examined to study the details of broken fibers. Furthermore, by using a marking

agent prior to the pyrolysis, delaminations are clearly visible and matrix cracks in the adjacent ply may also be

visible. This procedure is well suited to general angle-ply laminates of graphite/epoxy. The procedure for

deplying graphite/epoxy composites was originally developed by S.M. Freeman (Ref 89).

Footnote

The section “Interlaminar Shear Properties of Fiber-

Reinforced Composites at High Strain Rates” was written

by John Harding and Stephen Hallett, Oxford University. The section “Fatigue Testing and Behavior of Fiber-

Reinforced C

omposites” was written by W. Steven Johnson and Ramesh Talreja, Georgia Institute of

Technology.

References cited in this section

85. Manual on Experimental Methods for Mechanical Testing of Composites, R.L. Pendleton and M.E.

Tuttle, Ed., Society for Experimental Mechanics, Galione, 1989, p 170

86. G.A. Etheridge, W.S. Johnson, and S. Reeve, Effect of Lay-Up and Constraint on Tensile Notch

Strength, Proc. of the American Society of Composites 13th Annual Technical Conf. on Composite

Materials, A.J. Vizzini, Ed.,1998, p 489–504

87. J.G. Bakuckas Jr., W.H. Prosser, and W.S. Johnson, Monitoring Damage Growth in Titanium Matrix

Composites Using Acoustic Emission, J. Compos. Mater., Vol 28 (No. 24), 1994, p 305–328

88. W.S. Johnson, Screening Metal Matrix Composites Using Ultrasonic C-SCANS, ASTM J. Compos.

Technol. Res., Vol 11 (No. 1), Spring 1989, p 31–34

89. S.M. Freeman, Characterization of Lamina and Interlaminar Damage in Graphite-Epoxy Composites by

Deply Technique, ASTM STP 787, 1982, p 50–62

Mechanical Testing of Fiber-Reinforced Composites

Dale Wilson, The Johns Hopkins University, Leif A. Carlsson, Florida Atlantic University

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