Journal of Composite Materials 2005 Berkowitz 1417 31

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http://jcm.sagepub.com/content/39/16/1417The online version of this article can be found at:

DOI: 10.1177/0021998305050432

2005 39: 1417Journal of Composite MaterialsC. Kyle Berkowitz and W. Steven Johnson

Fracture and Fatigue Tests and Analysis of Composite Sandwich Structure

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Fracture and Fatigue Tests and Analysisof Composite Sandwich Structure

C. KYLE BERKOWITZ*

G. W. Woodruff School of Mechanical Engineering

Georgia Institute of Technology

Atlanta, GA 30332, USA

W. STEVEN JOHNSON

School of Materials Science & Engineering and

G. W. Woodruff School of Mechanical Engineering

771 Ferst Dr., Georgia Institute of Technology

Atlanta, GA 30332-0245, USA

(Received March 26, 2004)(Accepted October 13, 2004)

ABSTRACT: A composite sandwich system is investigated in this research. Quasi-static fracture toughness and fatigue crack growth experimental and analyticalapproaches are the focus. The particular system studied is comprised of a Nomex

(aramid fiber) honeycomb core with graphite/epoxy facesheets (skins).A modified version of the double cantilever beam (DCB) specimen geometry is

used for experimentation. The critical strain energy release rate, Gc, is used to

characterize the fracture toughness of the facesheetcore joint. Fatigue crack growthtesting is also performed. Novel analytical and experimental techniques are coupledand utilized to address challenges presented by the material system, especiallydifficult crack visualization. Crack length and growth can be estimated with anempirical approach, employing a compliance calibration. Experiments can also besimulated once several constants are estimated, aiding design. Many of thesetechniques can be generalized to other adhesive DCB experimentation.

Results show that cold tests result in higher fracture toughnesses and slightlyslower fatigue crack growth rates than room temperature tests. The hot temperaturehas less significant impact. Although only a limited amount of very slow growth data(


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INTRODUCTION

SANDWICH STRUCTURES ARE normally composed of high strength thin skins

(facesheets) bonded to a significantly thicker, less dense, and weaker core material.

Sandwich structures are most useful in applications that put a premium on stiffness-

to-weight ratios, such as in many aerospace applications. The focus of this paper is

a sandwich structure with graphite/epoxy facesheets and a Nomex (aramid fiber)

honeycomb core. This and similar systems appear on many commercial jets, especially as

control surfaces. These composite structures tend to develop damage through long-term

use, and this creates an enormous maintenance concern since the structural properties

are strongly dependent on the bond integrity. These debonds can occur in the film

adhesive used between the facesheets and core, or damage can occur in the core itself. The

goal of this research program is to characterize the fracture and fatigue parameters of a

particular material system, so that they can be incorporated into broader life predictions

and a damage tolerant methodology. The properties were experimentally measured, as

functions of test temperature, and novel experimental and data analysis strategies wereemphasized.

BACKGROUND

Sandwich structures have been widely applied in both aerospace and marine

applications, as well as anywhere where weight and stiffness are the driving performance

parameters. The first production parts that employed the idea of the modern sandwich

structure were produced in the 1940s, but the basic idea existed much earlier [1]. While

sandwich structures have been widely researched, only in the past two decades hasthis work addressed fracture and failure issues. Most of this work has focused on

structures with foam core materials, and it has addressed both experimental fracture

toughness determination and FEA to model failure criteria and predict parametric

dependencies [2,3]. Sandwich buckling and post-buckling analysis has received much

attention as well.

Avery and Sankar [4] and Ratcliffe and Cantwell [5,6] are among the only researchers in

the open literature who have performed experimental fracture work on Nomex-cored

sandwich structures, which is a very common material in commercial aviation. Fatigue

debond growth, as in the classic Paris formulations used in metals, has virtually been

ignored or not well reported. The maintenance of commercial aircraft, which are used

at a very high capacity for decades in many cases, seems to dictate a need for better

characterization of structures of this type.

Nomex aramid mechanical paper is a synthetic material composed of aramid fibers

stabilized with a phenolic resin. Sheets are then expanded and bonded together in the

honeycomb shape. Nomex has been shown to be very resistant to heat and moisture [7,8].

Polymers properties are generally a function of temperature; these effects are often

undesirable for an application. Aerospace composites are usually exposed to large

temperature ranges that significantly affect their mechanical behavior. Johnson and

Butkus [9] examined the compound effects of temperature and environmental aging on

aerospace-bonded joints. These conditions can affect both fracture toughness and fatigue

curves. Moreover, it is important to examine how environmental factors, especiallytemperature, affect aerospace composites.

1418 C. K. BERKOWITZ ANDW. S. JOHNSON

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EXPERIMENTAL PROCEDURES

Materials and Specimen Preparation

The specimens used in this research were constructed from six layers (plies) of woven

carbon fiber-epoxy prepreg, all oriented squarely to the rectangular specimens. Each

six-ply facesheet had a total cured thickness of1.5 mm. A thin (relative to a ply) film

adhesive was placed between the prepreg and core, and it was co-cured with the prepreg.

The honeycomb core had a cell size of 3.2 mm (1/8 in.), nominal density of 48 kg/m3

(3.0 lb/ft3), and thickness of 9.5 mm (3/8 in.). The prepreg and honeycomb were Hexcel

products W3F282-T6000 -F593 and HRH-10-1/8-3.0, respectively. The adhesive was Cytec

Metlbond 1515-3M.

The specimens were made from the manufactured rectangular panels, which had the

approximate dimensions of 610 mm 483 mm 11.9 mm (24 in. 19 in. 0.47 in.). The

panels were constructed from their layers of constituents and vacuum bagged, before being

cured in an autoclave at 177C (350F) for 2 h at an absolute pressure of 4 atm. The co-curing of the facesheets and adhesive film created a blended interface, where there was

not a clear boundary between the two. Their viscosities during curing allowed a certain

amount of mixture to occur. Since the facesheets and adhesive were cured as they were

being bonded to the core under the autoclaves pressure, the core became partially

embedded in the facesheet/adhesive layer. This manufacturing technique likely leads to a

stronger joint and certainly affects interface properties.

Each specimen, which was tested in a double cantilever beam (DCB) configuration, was

cut from a panel to a size of 152 mm 102 mm 11.9 mm (6 in. 4in. 0.47 in.), with the

ribbon direction of the honeycomb aligning with the longer dimension; the 102 mm (4 in.)

width was necessary to increase the loads to a reasonable level for the load cell. Pianohinges were riveted to the specimen to facilitate loading, and one edge was painted white to

ease visual crack measurement. A starter crack was cut with a saw between one facesheet

and the core, 20 mm past the load line; the crack grew in the ribbon direction. A top view

photograph of a specimen is shown in Figure 1 (upper half of hinge is not shown to reveal

the rivets), and a side view sketch is shown in Figure 2.

Mechanical Testing

Identical specimens were used for the fracture and fatigue testing; only the loading and

data recording schemes differed. A servo-hydraulic universal mechanical testing machine

was used, and mechanical wedge grips were used to grasp the free halves of the hinges

used to load the specimens. Since no standards existed for fracture toughness testing

of sandwich structures, ASTM D5528-94a was used only as a guideline when appropriate

[10]. The specimens were loaded via the piano hinges at a constant displacement rate of

1.27 mm/min (0.05 in./min). Loads and displacements were recorded by the testing

software, and visual crack measurements were taken. The actuator was programmed to

displace 0.5 mm (0.020 in.) past the point of nonlinearity, which produced crack growth.

Then, the specimen was unloaded, and the test was repeated 1520 times per specimen,

each giving a fracture toughness value.

Fatigue testing was performed at 4 Hz and at anR-ratio (Pmin/Pmax) of 0.1. The goalwas to obtain da/dNversus G curves for the parameters studied. Displacement control

Fracture and Fatigue Tests and Analysis of Composite Sandwich Structure 1419



5/16

was used because the loads were relatively small for control feedback. The load was

allowed to drop 10% before the test was stopped and restarted. At the beginning and end

of each segment of cycles, peak loads and compliance were measured for use in analysis,

and periodic visual measurements were taken.

Testing was performed at hot (77C (170F)), room (21C (70F)), and cold (54C

(65F)) temperatures, mimicking the extremes a commercial aircraft should see in

service. The temperature-controlled environment was created with an insulated chamber

mounted to the test machine, having both resistance heater and liquid nitrogen capabilities

and an electronic controller. Photographs of a specimen in the testing machine are shown

in Figure 3.

ANALYSIS

Fracture Toughness

The critical strain energy release rate, Gc, was used to characterize fracture toughness

according to:

G P2

2BdCda

, 1

Figure 1. Photograph of DCB specimen (top view).

Core

Piano hinges Debond

P

P

11.9mm (0.47in)Facesheets

152mm (6.0in)

Figure 2. Sketch of DCB specimen (side view).




6/16

whereP is the applied load,B is the specimen width (102 mm),Cis the compliance, anda

is the crack length. The critical, nonlinear load (Pc) corresponds to Gc. Equation (1)

characterizes the energy consumed to create an increment of new crack area and applies to

most cases of LEFM with constant specimen width. In this case, nominal area was taken;

no consideration was given to the geometry of the honeycomb.

Since no exact solution exists for this geometry due to asymmetry and the cores highorthotropy, dC/da was computed from a very good power law correlation that resulted

from compliance versus crack length data for a number of specimens. A plot of the data

used to develop this correlation is shown in Figure 4. This technique is often called the

Berry Method [11]. This correlation is expressed as:

C Dam, 2

where D and m are fitting parameters. An equation of this form is easily differentiable

for use in Equation (1). The correlation for the specimens used here was found to be:

C 2:01 106a2:69, 3

where C is in units of mm/N and a in mm. Equation (2) can be inverted, such that a

measured compliance can estimate crack length. Utilizing that fact, Equations (1) and (2)

can be combined as:

G

P2

2B mD

C

D 1=m" #

m1

, 4

Figure 3. DCB specimen loaded in test machine.




7/16

which reduces G to a function of machine-measured parameters and known constants.

This implies that the manual measurement of crack length is not needed once Equation (3)

is known, since compliance and crack length are 1-to-1. This assumption is validated by

the quality of the simple model of Equation (2) in fitting the actual visual measurements.

Many other models exist to predict and correlate compliance, but a more complex model

was not warranted here.

It may be unusual to use measured compliances in lieu of measured cracks in determining

applied G. It is usually preferable to apply Equation (2) with compliance as the dependant

parameter. However, for these particular experiments, it was found that achievingrepeatable visual measurements of crack lengths and crack extensions (a) was difficult

due to the mode of failure, specimen width, and experimental apparatus. The fracture line

of the crack was difficult to characterize, although an approximate crack length was always

obtainable, allowing the compliance correlation to be developed from many specimens.

Sensitivity and repeatability are important experimental issues with fracture and fatigue

testing, due to the strong power dependencies of the resulting properties. Since compliance

measurement was found to have much higher relative resolution than crack measurement,

compliance was preferred as the input to data reduction. Moreover, compliance was much

more repeatable and reliable for use inG calculations, whereas visual crack measurements

seemed to cause unnatural scatter in the data, again resulting from the visualization issues.Therefore, the numerical compliance calibration approach described here can be a useful

alternative in certain experimental circumstances.

Fatigue Crack Growth

CRACK GROWTH RATE CALCULATIONS

Fatigue crack growth per cycle (da/dN) can be characterized by a version of Paris

equation [12]:

dadN

c G n, 5

Compliance vs. crack length: calibration curve

C = (2.01 x 106) a2.69

R2= 0.965

0

0.1

0.2

0.3

0.4

0.5

10 25 40 55 70 85 100 115

Crack length (mm)

Compliance(mm

/N)

Composite visual data

from 21 DCB fracture

toughness specimens

Figure 4. Compliance calibration.




8/16

where c and n are fitting parameters, and G Gmax Gmin. In this research G,

expanded in Equation (6), was approximately 99% ofGmaxbecause anR-ratio (Pmin/Pmax)

of 0.1 was used: (Pmax)2 (0.1 Pmax)

2 0.99 Pmax. Therefore, there is no significant

distinction in this case between G and Gmax, as shown by:

G P2max P

2min

2B

dC

da Gmax: 6

Evaluating G utilizing the compliance calibration technique was performed identically as

in the monotonic fracture testing, and the applied G was calculated using Equation (6).

Pmax and Pmin were obviously used instead of the critical nonlinear Pc, which results in

quasi-static fracture failure.

Perhaps more useful than using compliance to correlate with crack length is using

compliance changes to correlate with crack growth. This is particularly valuable when

visual measurements of cracks are difficult, and it also allows for greater automation. The

use of compliance changes to calculate da/dNis fairly straightforward. Compliance was

measured frequently during the fatigue testing to ensure that incremental increases were

captured. Compliance was measured at monotonic test speeds, because this allowed for

much greater resolution. Loads were chosen to easily avoid Pc

for a given condition.

Small increases in compliance allows for estimation of small changes in crack length to

be performed. The measuring of small changes in crack length is important because it

generates more data per specimen and lessens load drop effects, which are described later

in this paper. Equation (7) can be used to express a change in crack length during a

segment of cycles, from cycle N1 to N2, in terms of compliance increase from C1 to C2.

a12 C2=D1=m C1=D

1=m, 7

whereD andm are defined in Equation (2). da/dN is simply evaluated with Equation (7)

and using the machine count of cycles, N N2 N1.

It is worth noting that this technique estimates a change in crack length from a change

in compliance, not the crack length itself from compliance measurements as in the

monotonic case. This fact decreases the sensitivity ofda/dNto possible crack measurement

errors from the compliance calibration expression. There are many potential numerically

more sophisticated ways of performing these calculations, although they were not

necessary for the work presented here. Numerical differentiation for these purposes are

further documented in ASTM E647-00 [13].

LOAD SHEDDING EFFECTS AND EXPERIMENTAL STRATEGY

As displacement-controlled testing results in load shedding, care must be taken to

ensure that the applied range ofG is approximately constant for eachda/dNdata point.G

drops dramatically with load, so frequent monitoring of load and compliance is prudent;

however, it is not clear how many cycles should pass between each measurement of a

da/dNdata point. If this is performed too often, machine noise could become noticeable

and interfere with the desired material property. If this is performed too infrequently, G

might reduce dramatically during segments of cycles used in da/dN calculations, which

would make the data difficult to interpret. The Gat the beginning and ending of a segmentof crack should be calculated and the consequences of assumptions considered.




9/16

There is some question as to how much the G should be allowed to drop in a

displacement-controlled test before taking measurements. It is impractical to take

measurements constantly, which would be time consuming and create noise in da/dN

data. Obviously, a resolvable growth in crack length must be allowed to take place over

time as G falls. It is best to understand the falling driving force behavior and

incorporate it into the experimental strategy. The effect of changes in G can be

thought of on a loglog plot of Equation (5), which is linear with slope n. A change in

log (G) is equal to log {(G1/G2)}, where G1 occurs at the beginning of a cycle

segment and G2 after load is shed. If n is assumed to be 4 (this can be checked after

some data is measured), which is reasonable for many composites [9], a value of 0.0833

for log {(G1/G2)} would result in 1/3 decades of change in log (da/dN). This seems

somewhat reasonable, and scatter bands are often this large for general da/dN

data. Therefore, this amount of drop in G was tolerated before taking measurements.

Log (P21 max=P22 max) is approximately proportional to log {(G1/G2)}, so this was

allowed to reach 0.0833 before each segment of cycles was stopped for measurement; i.e.,

(P1max /P2max) 1.10.Additionally, a minimum compliance increase of 5% (C2/C1 1.05) was prescribed to

occur between cycle counts used to calculateda/dN. This prevented excessive noise in the

data by ensuring a resolvable crack extension had occurred. All data points were checked

against these criteria. The load drop and compliance increase restriction imply there is

only a window of cycles during which accurate data can be gathered. If approximate

growth rates can be estimated, specific cycle counts between necessary measurements can

be approximated.

The largest (beginning) Gwas associated with measured da/dNgrowth for a given set

of cycles. This may seem slightly nonconservative, but it is significantly more accurate

than averaging because of the power nature of the Paris Law. Conservatism can be addedto the data after it is accurately calculated by adding simple safety factors. In this

case, shifting the da/dNdata 1/3 decade upward may make sense to create a worst-case

boundary for the measured data.

Due to the unusual nature of the analytical approach presented here, an examination of

the results of a hypothetical test may help evaluate the method for a particular use. If the

compliance calibration and crack growth empirical constants, as shown in Equations (2)

and (5), can be estimated, then the researcher can simulate the fatigue crack growth

process and following data reduction and analysis. This can be done to aid in specimen

design and to refine experimental strategy before an expensive and time consuming

experimental fatigue crack growth task is begun. For instance, one common pitfall is an

excessively compliant specimen, which can result in very high deflection not conducive

to high-frequency fatigue testing. A fatigue simulation using the empirically derived

constants measured for this research is shown in Table 1. Figures 5 and 6 aid the

visualization of several parameters throughout the experiment.

The simulation is of a single specimen, loaded in a sequence of 3 blocks. Each of these

blocks of fatigue cycles begins at a chosen load that produces reasonably high crack

growth rates. Then, as load is shed, the growth rates near an assumed threshold rate. Then

the test is reset, increasing the load (and displacement) for another period of crack

growth. Several snapshots in time are shown during the fatigue test with associated

parameters, including cycle count (time). The compliance and da/dN are calculated via

the assumed empirical constants, and the other parameters are derived according tothe previously mentioned analysis. Note the total time would be 57 days at 5Hz.




10/16

Table 1. Simulation of DCB fatigue crack growth test.

Block Pmax(N) max(mm) C (mm/N) a (mm) dC/da Gmax (J/m2) da/dN N (cycles) log (N)

1 350 2.28 0.00651 20.0 0.00088 530 8.4E-04 1 0.0

1 280 2.28 0.00814 21.7 0.00101 390 3.1E-04 5487 3.7

1 224 2.28 0.01018 23.6 0.00116 288 1.2E-04 21,340 4.31 179 2.28 0.01272 25.6 0.00134 212 4.4E-05 67,153 4.8

1 143 2.28 0.01590 27.8 0.00154 156 1.7E-05 199,539 5.3

1 115 2.28 0.01988 30.2 0.00178 115 6.3E-06 582,102 5.8

1 92 2.28 0.02485 32.8 0.00204 85 2.4E-06 1,687,616 6.2

1 73 2.28 0.03106 35.7 0.00235 62 8.9E-07 4,882,280 6.7

2 190 5.90 0.03106 35.7 0.00235 418 3.9E-04 4,882,280 6.7

2 152 5.90 0.03882 38.7 0.00271 308 1.5E-04 4,903,257 6.7

2 122 5.90 0.04853 42.1 0.00311 227 5.5E-05 4,963,874 6.7

2 97 5.90 0.06066 45.7 0.00358 167 2.1E-05 5,139,043 6.7

2 78 5.90 0.07582 49.6 0.00412 123 7.8E-06 5,645,239 6.8

2 62 5.90 0.09478 53.9 0.00475 91 2.9E-06 7,108,020 6.9

2 50 5.90 0.11847 58.6 0.00546 67 1.1E-06 11,335,099 7.1

3 120 14.22 0.11847 58.6 0.00546 387 3.1E-04 11,335,099 7.1

3 96 14.22 0.14809 63.6 0.00629 285 1.1E-04 11,379,035 7.1

3 77 14.22 0.18511 69.1 0.00723 210 4.3E-05 11,506,000 7.1

3 61 14.22 0.23139 75.0 0.00833 155 1.6E-05 11,872,899 7.1

3 49 14.22 0.28924 81.5 0.00958 114 6.1E-06 12,933,147 7.1

3 39 14.22 0.36155 88.5 0.01103 84 2.3E-06 15,997,004 7.2

3 31 14.22 0.45193 96.2 0.01269 62 8.6E-07 24,850,797 7.4

Load, strain energy release rate, and displacement vs. cycles

0

100

200

300

400

500

0.0E+0 2.0E+6 4.0E+6 6.0E+6 8.0E+6 1.0E+7 1.2E+7 1.4E+7 1.6E+7 1.8E+7 2.0E+7

N (cycles)

LoadandG(lbandJ/m2)

0

3

6

9

12

15

(mm) P

G

d

Figure 5. Load, strain energy release rate, and displacement during fatigue simulation.

Crack length and crack growth rate vs. cycles

0

30

60

90

120

0.0E+0 2.0E+6 4.0E+6 6.0E+6 8.0E+6 1.0E+7 1.2E+7 1.4E+7 1.6E+7 1.8E+7 2.0E+7

N (cycles)

Cracklength(mm)

1.0E-07

1.0E-06

1.0E-05

1.0E-04

1.0E-03

Crackgrowthrate

(mm/cycle)

av.N

da/dN v. N

Figure 6. Crack and crack growth rate during fatigue simulation.




11/16

A researcher could use this type of simulation to examine trade-offs among priorities and

potential problems, and their causes.

RESULTS

The test matrix of the experiments performed is shown in Table 2. Eleven quasi-static

(monotonic) fracture toughness tests were performed and seven fatigue tests. The

experiments were performed at three different test temperatures: hot, cold, and room

temperatures, at the combinations described in the matrix.

Fracture Toughness

The numerous Gc measurements from each test condition were averaged, with eachspecimen being weighted equally, and they are shown in Table 3. Many fracture toughness

measurements for each specimen were obtained using the loadcrackunload technique

previously described. A typical specimens loaddisplacement plot obtained for four load

unload cycles is shown in Figure 7.

Many materials exhibit a crack length dependence on Gc. This is generally attributed to

a change in plastic zone shape with crack growth [14]. However, no dependence on crack

length was found for fracture toughness for any condition studied in this work. This

suggests that the failure can be more easily characterized by the single parameter of strain

energy release rate. The lack of crack length dependence is often referred to as a flat

R-curve. A plot ofGc versus crack length is shown in Figure 8.Clearly, the test temperature had an effect on fracture toughness. The cold temperature

tests resulted in the highest fracture toughness, and the hot temperature tests had the

Table 2. Test matrix.

Fracture

toughness Fatigue

Test temperature (monotonic) crack growth

Room temperature (21C) 6 2

Cold temperature (

54C) 3 2

Hot temperature (77C) 2 3

Total 18

Table 3. Fracture toughness results.

Specimens Values

Gc,avg(J/m2)

Room temperature (21C) 6 56 1180

Cold temperature (54

C) 3 62 1620Hot temperature (77C) 2 42 1160




12/16


13/16

ribbons of the honeycomb. There was likely not a homogeneous plastic zone, and the

failure could have been more dominated by the aramid fibers strength than by any stressconcentrating effect within the Nomex material; however, this idea would need further

validation. This may be an explanation for the increase in fracture toughness at the

cold temperature; for colder temperatures generally increase a materials strength, while

generally decreasing fracture toughness.

In most applications, the core failure mode would be desirable and designed, since that

would imply a high quality bond at the adhesive interface, which is important to a

sandwich structures integrity. It is worth noting that the mode of failure and fracture

toughness values can be strongly dependent on processing, in addition to the material

constituents. For instance, another sandwich structure of identical facesheets, core, and

adhesive as the one studied here, except with precured facesheets secondarily bonded to

the core, could have considerably different failure behavior than a co-cured structure. The

presence (or not) of an adhesive film may also have a large effect; and clearly the curing

cycle, tooling, and constituents themselves are important parameters. Due to these

considerations, little can be generalized for other composite sandwich structures based on

these results. In fact, these issues warrant further research.

Fatigue Crack Growth

The fatigue crack growth results are presented by temperature in Figure 10. The

resulting trends from the fatigue data are consistent with those found from the quasi-staticfracture toughness tests, and the mode of failure was the same as in the fracture case.

Figure 9. DCB fracture surface.




14/16

The linear shape of the data on Figure 10 verifies the validity of the Paris crack growth

model that was applied. The data agrees with the model well in the range of crack growth

rates between 5 106 and 5 103 mm/cycle. A line has been sketched on the figure

showing that an approximate Paris exponent of 3.2 fits the data well in this regime. This

line is not intended to fit any particular set of data, but is added for visual purposes.

A few slower growth rate data points suggest the Paris relationship is not at all

valid below 1 106 mm/cycle, which is quite normal for many materials as threshold

is approached. Although a true threshold was not verified, a somewhat arbitrary con-

vention for composite fatigue is to assume 1 106 mm/cycle is significant, and a line

Fatigue crack growth rate vs. strain energy release rate

1.0E-07

1.0E-06

1.0E-05

1.0E-04

1.0E-03

1.0E-02

10 100 1000 10000

G (J/m2)

da/dN(mm/cycle)

RT (21C)

CT (54C)

HT (77C)

Approx. Paris fit

Freq=4 Hz

R=0.1

da/dN=1.6E-12 (G)3.2

Threshold

GIC

Figure 10. Fatigue crack growth data.

Fatigue crack growth rate vs. strain energy release rate

1.0E-06

1.0E-05

1.0E-04

1.0E-03

100 1000

G (J/m2)

da/dN(mm/cycle)

RT (21C)

CT (54C)

HT (77C)

Freq=4 Hz

R=0.1

Figure 11. Fatigue crack growth data zoomed-in.




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corresponding to this is sketched on the figure at 70 J/m2. The data shown in this slow

growth regime supports that a threshold effect exists. Another line at 1200 J/m2 is shown

on the figure; this is an approximation for fracture toughness.

Figure 11 shows a zoomed-in view of the data in Figure 10. In this plot, the data points

can be more easily distinguished and compared. Although the overlap of the data shown in

the figure lessens the certainty with which conclusions can be drawn, the cold temperature

resulted in slower crack growth rates than room temperature. This would be expected

since the cold temperature increased toughness. As shown in the figure, the hot

temperature data seems to be highly overlapped with the room data. The point averages

would suggest that the hot temperature resulted in faster growth, but the lack of

separation of the scatter bands on the figure causes any conclusions to be unclear. Similar

to the fracture toughness results, the cold temperature had a more obvious effect on the

fatigue results than the hot temperature, relative to the baseline of room temperature.

CONCLUSIONS

A fracture mechanics-based approach has been utilized for characterizing a sandwich

structure common to commercial aviation. Fracture toughness and fatigue crack growth

testing of a particular sandwich system were performed on double cantilever beam style

specimens. These specimens were tested at each of the three temperatures: hot (77C),

room (21C), and cold (54C). Compliance versus crack length correlating was shown as

an effective tool in deducing both applied strain energy release rates and crack growth

rates from test data. Using compliance to estimate crack length can be a useful technique

when a high resolution in compliance change with crack extension is achievable, especially

in cases where crack tip visualization is problematic. When the necessary empiricalconstants can be estimated, a fatigue test can be simulated. This can be a valuable tool for

the design of specimens and experimental strategy when a compliance crack length

calibration can be utilized.

The temperature effects on the particular sandwich system studied were significant. The

cold temperature increased toughness and reduced fatigue crack growth rates, relative

to room temperature. The hot temperature had relatively little impact; however, any

marginal effects were the opposite trends as those caused by the cold. A Paris crack growth

model fit the data well for most growth rates between near-threshold and near-Gc, and

a flat R-curve behavior was exhibited by the material.

The failure of all specimens tested was in the core material. This suggests that coreproperties were being measured. The mode of failure is possibly due to manufacturing

techniques, as much as the constituents. The core was weaker than the bonding of the

facesheets to the core. The mechanisms of failure and resulting data suggest that core

strength is the dominant factor in the fracture mechanics properties presented in this

paper.

REFERENCES

1. Zenkert, D. (ed.) (1997). The Handbook of Sandwich Construction, EMAS Ltd, UK.

2. Falk, L. (1994). Foam Core Sandwich Panels with Interface Disbonds, Composite Structures,28(4): 481490.




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3. Carlsson, L.A. and Prasad, S. (1993). Interfacial Fracture of Sandwich Beams, EngineeringFracture Mechanics, 44(4): 581590.

4. Avery, J.L. and Sankar, B.V. (2000). Compressive Failure of Sandwich Beams with DebondedFace-Sheets, Journal of Composite Materials, 34(14): 11761199.

5. Ratcliffe, J. and Cantwell, W.J. (2001). Center Notch Flexure Sandwich Geometry for

Characterizing Skin-Core Adhesion in Thin-Skinned Sandwich Structures, Journal of ReinforcedPlastics and Composites, 20(11): 945970.

6. Ratcliffe, J. and Cantwell, W.J. (2000). A New Test Geometry for Characterizing Skin-CoreAdhesion in Thin-Skinned Sandwich Structures, Journal of Materials Science Letters, 19(15):13651367.

7. Hentschel, R.A.A. (1975). Nomex Aramid PapersProperties and Uses, In: TAPPI PaperSynthetics Conference, Atlanta, GA, October 68, pp. 189213.

8. Shafizadeh, J.E. and Seferis, J.C. (2000). The Effects of Long Time Water Exposure on theDurability of Honeycomb Cores, In: 45th International SAMPE Symposium and Exhibition,Covina, CA, May 2125, pp. 380388.

9. Johnson, W.S. and Butkus, L.M. (1998). Considering Environmental Conditions in the Designof Bonded Structures: A Fracture Mechanics Approach, Fatigue and Fracture of Engineering

Materials & Structures, 21(4): 465478.10. ASTM D5528-94a (1994). Standard Test Method for Mode I Interlaminar Fracture Toughness

of Unidirectional Fiber-Reinforced Polymer Matrix Composites, In: Annual Book of ASTMStandards, American Society for Testing and Materials, Philadelphia, PA.

11. Berry, J.P. (1963). Determination of Fracture Energies by the Cleavage Technique,Journal ofApplied Physics, 34(1): 6268.

12. Paris, P.C. and Erdogan, F. (1963). A Critical Analysis of Crack Propagation Laws,Journal ofBasic Engineering, 85: 528534.

13. ASTM E647-00 (2000). Standard Test Method for Measurement of Fatigue Crack GrowthRates, In: Annual Book of ASTM Standards, American Society for Testing and Materials,Philadelphia, PA.

14. Anderson, T.L. (1995).Fracture Mechanics: Fundamentals and Applications,2nd edn, CRC PressLLC, Boca Raton, FL.


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