fatigue performance and analysis of composite patch ... et al_fatigue performance of composite...

Paper No. Year-2014 Hart 1

Fatigue Performance and Analysis of Composite Patch Repaired

Cracked Aluminum Plates

Author Name(s): D. C. Hart

1 (V), Udinski, E. P.

1 (V), Hayden, M. J.

1 (V), Dr. Liu, X.

2 (V)

E-Glass epoxy composite patches offer repair method for cracks in sensitized aluminum. Patches

are water tight, low modulus, mitigate crack growth, and eliminate “hot work”. Patched CCT

specimens were tested in tension-tension fatigue with 0.25 inch aluminum as-delivered and oven-

sensitized plate. Composite patch repairs increased fatigue cycles to failure by more than 4 times.

1 Naval Surface Warfare Center, Carderock Division, Survivability, Structures and Materials Department

2 Global Engineering and Materials, Inc.

KEY WORDS: Composite; Center Crack Tension, Middle

Tension, Patch, Sensitization.

INTRODUCTION

Research, testing, and analysis presented in this paper

provides proof of concept and documentation of the effect

composite patch repair has on the response of center crack

tension (CCT) specimens due to static and fatigue loads. Large

scale CCT testing and analysis supports the development of the

composite patch repair method in the US Navy, which is

currently experimenting with alternative repair methods to

address sensitized aluminum superstructure cracking. Aluminum

alloys containing magnesium are susceptible to sensitization and

subsequent intergranular corrosion(O'Shaugnessy, Nov 1985),

possibly making the cracked plate un-weldable.

One alternative to traditional welded repair is the

composite patch. Composite patch repairs have seen significant

use by the aerospace community on thin aluminum aircraft skins

and some thicker cast components while only limited use on

marine structures to address stress related cracking(Grabovac,

2003) and corrosion of steel plating(Weitzenbock, 2012). To

date, composite patch repair prototypes have been installed on

10 ships and have demonstrated their ability to inhibit water

ingress and to maintain a weather tight compartment. At the

time of this report 1300 sq ft of composite repair patches have

been successfully installed and in operation with the first

installation performed in December of 2010.

Deck repair using a composite patch offers a significant

repair cost and safety benefit from the fact that applying the

patch does not require “hot work”. Hot work refers to torch

cutting and welding tasks that require a fire watch and a clear

area on both sides of the repair and when in the vicinity of a

flammable substance requires the spaces to be gas free.

The strength and fatigue life benefits of the composite

patch repair method have been well documented. Static testing

of edge notch panels repaired with a composite patch

demonstrated a significant increase in panel strength with a thin

composite patch across the crack tip. Testing results showed that

a single layer of unidirectional E-Glass increased ultimate

strength of 0.25 inch thick specimens by 33% and 37% for

0.125 inch thick aluminum plate (Bone, 2003).

Fatigue performance benefits were demonstrated for three

configurations of 3-ply unidirectional boron-patched, stiffened

0.04 in. thick aluminum plates were tested to study the tension-

tension fatigue crack growth rate: cracked plate, cracked

stiffened plate with 4 in. spacing, and cracked stiffened plate

with 6 in. spacing. FEA was utilized to calculate the Stress

Intensity Factor (K) and modeled the crack growth using Paris

Law coefficients for the underlying plate. Test and analysis

results showed that the composite patches decreased the crack

growth rate and increased the fatigue life of the repaired

panels(Sabelkin, 2006).

Previous work was done testing 0.25 inch thick aluminum

center crack tension specimens with stiffness-matched bonded

unidirectional boron patches to study effects of a composite

patch on the aluminum plate fatigue life. Results showed no

benefit of increased patch length and the composite patch

disbond did not occur until the crack had grown to more than

65% of the plate width. Post mortem inspection revealed an

elliptical crack front which lagged by up to 0.4 inches from the

patched to un-patched surfaces. In addition, the crack growth

rate varied between the patched and unpatched sides causing the

crack font(Schubbe, 1999).

Numerical investigations and testing of center crack

tension specimens repaired with uni-directional E-Glass epoxy

patches showed that with minimal surface reinforcement the

fatigue life of the cracked plate could be increased, while adding

more than four plies to the repair did not significantly affect the

fatigue life of 0.25 inch aluminum panels(Hosseini-Toudeshki

H., 2007).

Addressing the current cracking of sensitized aluminum

plate required a composite patch repair that could address the

uncertainty in aluminum material condition and the extent of

cracking observed while returning water tight integrity to the

compromised compartment and mitigating or stopping crack

growth. The initial composite patch repair design was based on

Paper No. Year - 2014 Hart 2

Classical Lamination Theory (CLT), Rose’s Model closed form

solution(Baker, 2002), and simple Finite Element Analyses

(FEA) reducing the stress intensity at the aluminum crack tip by

more than 25% (Noland, 2013). An E-Glass laminate with

quasi-isotropic behavior was selected, which had a low modulus

of 1.7 Msi, avoided galvanic corrosion risk associated with

carbon, and had low cost readily available constituents. The low

modulus also aided in reducing the stress concentration

associated with applying reinforcement to the surface.

Testing of patched metallic specimens for the aerospace

industry typically involved thin aluminum plate with high

modulus uni-directional fibers applied perpendicular to the

crack. A majority of composite patch testing was performed on

thin plate, or plate thickness less than 0.25 inch. Though 0.25

inch plate is considered thick by aerospace standards, for marine

structures 0.25 inch plate is considered thin. Testing

documented in this report extends the scale of the typical test

specimens by both thickness and size of the specimen. Test

specimens were 11 inch wide, 0.25 inch thick aluminum plate

with a 5 inch initial crack. Test data from both patched and

unpatched specimens will be compared with crack growth

predictions made using the hybrid structure evaluation and

fatigue damage assessment (HYSEFDA) toolkit for ABAQUS

developed by Global Engineering and Materials, Inc. (GEM).

OBJECTIVES The goal of current research is to provide the proof of

concept demonstration and to document the effect composite

patch repairs have on the tension-tension fatigue crack growth

of 0.25 inch thick 5456-H116 aluminum CCT specimens. Test

responses under static and tension-tension fatigue loads were

compared for as-delivered aluminum plate, patched as-

delivered, and patched oven-sensitized aluminum plate and then

compare those responses with analytical predictions made using

HYSEFDA.

The objectives of the tension-tension fatigue CCT testing

are:

1. Measure the tension-tension fatigue cycles to

aluminum failure for baseline aluminum and

composite patch repaired CCT specimens.

2. Observe and document the behavior of the composite

patch repair laminate during aluminum crack

propagation.

3. Compare composite patch repair performance of as-

delivered and oven-sensitized aluminum plate.

4. Correlate the analytical and test response of CCT

specimens to tensile fatigue loading.

MATERIALS The aluminum plate selected for test specimens was 0.25

inch thick 5456-H116 plate. Oven sensitized plate was

conditioned at 250 ºF for 10 days resulting in a Degree of

Sensitization (DoS) of 47.3 mg/cm2; measured in accordance

with ASTM G67.

A rubber toughed epoxy resin system (M1002) from Pro-

Set was selected based on the manufacturer advertised bond

strength and toughness. The Pro-Set M1002 was mixed with

Pro-Set 237 hardener.

The laminate schedule followed used commercially

available E-Glass fabric either woven or stitched together. Ply

materials for the composite patch laminates include the 9.6

oz/yd2 Hexcel 7500 style 1:1 plain weave (PW) fabric, the 8.8

oz/yd2 Hexcel 7781 style 8 harness satin (HS) weave fabric, and

biaxial stitch bonded uni-directional fiber fabric with 12 oz/yd2

±45 (S45) and the 18 oz/yd2 0/90 (BX) fiber orientations.

TEST SPECIMENS Test specimens are an aluminum panel cut to size with a

machined notch that was fatigue pre-cracked under Stress

Intensity (K) control to develop a sharp crack tip. Pre-cracked

specimens are then prepared for bonding, which consists of

mechanical abrasion, acid etch and distilled water rinse, and a

silane treatment applied. When the silane treatment has cured

each ply is hand wet-out on the surface and the full laminate

stack is then consolidated under vacuum; Typical laminate

configuration shown in Figure 1. The applied vacuum level was

10 in-Hg applied approximately 1 hour after mixing with a

single ply of bleeder cloth. Laminate was 30 inches long with

3.5 inch tapers on the top and bottom edges and full thickness

on the cut edges. In this configuration the full thickness laminate

extended 11.5 inches from the crack in both directions.

Figure 1: Typical composite patch repair configuration.

Full thickness rectangular specimens were cut to nominal

dimensions of 11 inches wide by 37.5 inches long. Initial

machined notches were oriented in the transverse-longitudinal

(T-L) material direction, as defined in ASTM E399, with the

load applied in the materials longitudinal-transverse (L-T)

direction normal to the crack plane. Fatigue pre-cracking of the

2W

2a

σ

σ

Crack

Crack

3.5” 15”

0.15”


aluminum plate was done at a constant stress intensity factor (K)

of 10 ksi-in1/2

at a frequency of 3 Hz until the desired initial

crack length of 5 inches was achieved. Compliance coefficients

used in calculating crack length for K-control were provided by

Fracture Technology Associates (manufacturers of the K-control

test system) specifically for the tested configuration: C0 =

0.24061, C1 = 1.56371, C2 = 34.2863, C3 = 174.072, C4 =

272.8647, and C5 = 142.403. Bolted joints on the top and bottom

consisted of seven bolts with 1.375 inch pitch and 1.375 inch

edge distances, which were connected to the test frame with

threaded collars. The fixture and specimen configuration details

are shown in Figure 2 and listed in Table 1. The test

configuration and nominal specimen configuration are shown in

Figure 3.

Figure 2: Test specimen dimensions.

Table 1: Test Specimen Dimensions.

Table 2: Test Specimen Matrix.

Figure 3: Load frame with CCT test specimen.

TEST MATRIX Performance of the composite patch repair technique on

as-delivered and oven-sensitized plate was compared with

baseline as-delivered aluminum plate. Fatigue tests were

conducted using an MTS Model 809.25 axial torsion hydraulic

test system. Table 2 lists the specimen configuration, aluminum

condition, peak far-field stress applied, and specimen ID for

panels made for this study. Panels were studied in four groups;

the first group provided baseline cracked aluminum plate

behavior, the second group provides behavior of patched

aluminum plate behavior, the third group was patched oven-

sensitized plate, and the final group demonstrates the fatigue

performance of shipboard aged patched aluminum plates.

1.375”

2.75”

11.0”2a0

Ø 0.625”

1.375”

1.375”

~30”

37.5”

0.125”

0.125”

WJ σσ

Specimen Width [in] 11

Specimen Length [in] 37.5

Patch Length [in] ~30

Number of Bolts 7

Bolt Diameter [in] 0.625

Pitch [in] 1.375

Bolt Edge Distance [in] 1.375

Bolt Side Distance [in] 1.375

WaterJet Notch (WJ) [in] 4.9

Tab Thickness [in] 0.125

Tab Length [in] 2.75

Configuration Condition

Far-Field

Stress

(ksi) R-Ratio

Specimen

ID

Aluminum As Delivered 5.0 0.2 16

Aluminum w/Patch As Delivered 5.0 0.2 11

Aluminum w/Patch Oven Sensitized 5.0 0.2 17







Total 9


INSTRUMENTATION Each specimen was fitted with axial strain gauges aligned

with the load direction and a ring gauge to monitor the CMOD

at the center of the panel. On patched specimens the ring gauge

was fitted to the notch on the bare aluminum side because the

notch was covered on the patched side. Along with the initial

half crack length (a0), the CMOD measurement is used to

calculate specimen compliance, current crack length, and stress

intensity for the control software commanding the test frame.

The axial strain gauge configuration is shown in Figure 4.

Gauges are Micro Measurements CEA-06-250UW-350.

Aluminum specimens had 6 gauges, while composite patch

repaired specimens had an additional axial gauge bridging the

crack on the laminate at the center of the specimen.

Figure 4: Strain gauge configuration shown from the front side (composite side).

The CMOD measurements were performed using a John

A. Shepic Co. SFDG-30 ring gauge, as mounted in the specimen

in Figure 5. The ring gauge has a maximum opening

displacement of 0.3 inches with output from 0 to 10 volts.

Gauge calibrations were performed prior to each test.

Figure 5: Ring gauge mounted in machined notch of specimen.

TESTING PROCEDURES Specimens and testing procedure were guided by ASTM

D3039, ASTM E647, and ASTM E561-08. Instrumentation was

calibrated for the initial load segment for each static test

specimen. If static testing required multiple load segments, the

initial calibration was retained and the testing restarted. Fatigue

testing was run until aluminum panel failure, which was defined

as full width crack growth and separation of the aluminum plate

ligaments. Specimens were loaded until failure of the aluminum

plate, during the higher stress loads the composite also failed.

Fatigue specimens were loaded in tension with load amplitude

ratios (R) of 0.1 and 0.2 and load cycle frequencies of 1 and 2

Hz. The 5.0 ksi stress test series was run with R = 0.2 and a load

frequency of 2 Hz. Specimens were tested in a temperature

controlled lab with the as manufactured moisture content.

Relative Humidity (RH) was not explicitly controlled. The lab

temperature was 70°F with approximate RH of 70-80%.

CENTER CRACK TENSION TESTING Fatigue testing was performed at three stress levels. The

initial stress level of 5.0 ksi is based on the typical peak von

Mises stress upper level superstructure deck plating due to

primary hull bending loads. The stress level of 10.9 ksi was

above the 7.5 ksi aluminum fatigue design limit for

superstructure subjected to cyclic loading(ABS, 2009) and was

assumed to be within the linear response range for composite

patched specimens. Peak applied stress was 14.5 ksi,

corresponding with the peak far-field stress level achieved by

the aluminum specimens and assumed to be well into the non-

linear response range.

Data Acquisition (DAQ) captured the initial, middle, and

final load cycles in blocks of 1000 cycles. Load cycle counts

were determined when post processing the test data and

compared with MTS controller data. A load cycle was counted

when the peak and valley loads were within 1% of the

maximum and minimum set loads. The total cycle count

included load cycles performed between periods of active DAQ.

Cycles between DAQ sets were calculated using the stop and

start time difference between data sets and the loading rate.

Cycles were counted to the point of failure, which was defined

as the aluminum crack extending across the full width of the

plate. In most cases the composite patch repair continued to

carry load for several cycles beyond full aluminum plate failure.

When available, the CMOD data was used to determine the load

cycle at which the ligaments failed. Prior to failure, the CMOD

as a function of load cycle becomes severely non-linear

indicating an increased crack growth rate. From test

observations, crack extension in one of the aluminum ligaments

dominates the response. Figure 6 shows a typical CMOD versus

load cycle response (L) and the region of the response used to

determine ligament failure (R). The segment of the CMOD

response, highlighted, typically has two distinct areas that

identify stages of ligament failure. When crack extension

reaches a critical threshold in one of the ligaments, the amount

of fracture energy available results in plastic deformation and

rapid crack growth, noted by the slope at #1. Once the first

ligament fails and stress is redistributed, CMOD growth returns

to a steady state after #2. The load cycle at which the ligament is

assumed to have failed is located at the beginning of the steady

state response. Similar behavior is exhibited during failure of the

second ligament at #3 and #4. Again, ligament failure is

assumed to have occurred when the response returns to steady

Front (Composite Side): OddBack Side: Even

3(2)

0.50”On Center

0.50”On Center

Back to Back Gauges (Black)Composite Gauge (Green)

Top

0.50”On Center

5 (4)

1(0)

7


state near #4. Even though the aluminum is fully separated, the

composite patch continues to support the cyclic load.

Figure 6: CMOD data used to determine ligament failures.

TEST RESULTS The following results are for un-patched aluminum CCT

specimens 14, 15, and 16, which were tested in their as-

delivered condition. Specimens 10, 11, and 12 were composite

patch repaired aluminum with the aluminum in the as-delivered

condition. Specimens 17, 18, and 19 were composite patch

repaired aluminum plate that was oven sensitized after

developing the initial crack under fatigue loading.

5 ksi Far-Field Stress Level The 5 ksi far-field stress level testing of specimens 16,

11, and 17 was performed at a load cycle rate of 2 Hz from 3 to

14 kips, or a far-field stress range of 1.0 to 5.0 ksi. Strain gauge

data was offset by the first load segment’s zero load strain data.

Each subsequent set of strain data was offset using the same

zero data.

The un-patched aluminum specimen, 16, failed at a cycle

count of 54,628. Ligament failure occurred first on the 0/1 strain

gauge side followed by failure on the strain gauge 2/3 side. The

failure surface on the strain gauge 0/1 ligament of specimen 16

is shown in Figure 7. The center notch of the specimen is on the

right. The extent of the fatigue pre-crack is faint and not visible

in the photograph, but extends to the left about 0.1 inch. The

smooth region in the center of the thickness extending past the

pre-crack represents slow fatigue crack growth due to testing.

Stable ductile crack extension begins to dominate when half of

the ligament remains. The stable crack extension shifts shear

planes with about 1/3 of the ligament remaining and extends

until a ligament of about 0.25 inch remains before unstable final

fracture occurs.

Figure 7: Specimen 16 (un-patched, as-received aluminum) fracture surface on the strain gauge 0/1

ligament.

Aluminum failure occurred at cycle counts of 284,300

and 234,800 for composite patch repaired specimens 11 and 17,

respectively. The composite patch repairs did not fail and

continued to carry load across the cracked aluminum plate.

Specimen 17 loading continued for an additional 51,000 cycles.

Testing was stopped when laminate disbond was observed

approximately 0.5 inches across the width of the panel and

centered about the crack shown in Figure 8. Specimen 11 was

loaded for 3,000 additional cycles past aluminum plate failure

without signs of laminate disbond around the crack. Application

of the composite patch repair increased the fatigue life of the

cracked plates by 5.2 and 4.3 times for the as-delivered and

sensitized aluminum, respectively.

Figure 8: Specimen 17 post fatigue laminate disbond; red line marks extents.

10.9 ksi Far-Field Stress Level The 10.9 ksi far-field stress level testing of specimens 15,

10, and 19 was applied at a load cycle rate of 1 Hz from 3 to 30

kips, or a far-field stress range of 1.0 to 10.9 ksi. Strain gauge

data was offset by the first load segment’s zero load strain data.

Each subsequent set of strain data was offset using the same

zero data.

The un-patched specimen, 15, failed at a cycle count of

5,305. Ligament failure occurred first on the 2/3 strain gauge

side followed by failure on the strain gauge 0/1 side within 8

cycles. The failure surface of specimen 15 on the strain gauge

2/3 ligament is shown Figure 9. The center notch of the

specimen is on the left. The smooth fatigue pre-crack extends

approximately 0.25 inch to the right and distictly ends. Almost

immediately after the pre-crack, it appears that stable ductile

crack growth occurs until arrest about half way through the

ligament (fatigue striations can be seen in the photograph). The

crack appears to extend under fatigue until about 1/3 of the

ligament remains when extension becomes unstable until final

fracture occurs.

Figure 9: Specimen 15 (un-patched, as-received aluminum) fracture surface on the strain gauge 2/3

ligament.

0.000

0.005

0.010

0.015

0.020

0.025

0.030

0.035

0.040

0.045

0.050

0 5000 10000 15000 20000 25000 30000 35000 40000 45000

CM

OD

(in

)

Cycle

CMOD Data Recorded at Cycle Load Peaks

Specimen #10 (N=1 to 40,000)

0.025

0.030

0.035

0.040

0.045

0.050

38250 38500 38750 39000 39250 39500 39750 40000

CM

OD

(in

)

Cycle

CMOD Data During Fracture Recorded at Cycle Load Peaks

Specimen #10 (N=1 to 40,000)

Return to stead state as crack

grows in remaining ligament

Large increase in CMOD associated

with plastic deformation and crack

growth during ligament failure

Significant plastic deformation

as second ligament fails

1

2

4

3

#15

Unstable FractureFatigue Fracture Ductile Fracture


Aluminum failure occurred at cycle counts of 39,890 and

34,323 for composite patch repaired specimens 10 and 19,

respectively. The composite patch repairs did not fail

immediately after the aluminum failed and continued to carry

load across the separated aluminum plate. Specimen 10 loading

continued for an additional 563 cycles. The test was stopped

when the laminate disbond (approximately 1.0 inch high and

centered about the crack) was observed to fully extend across

the width of the panel, as shown in Figure 10. Figure 11 shows

the specimen 10 aluminum fracture surface for the strain gauge

0/1 ligament. The center notch of the specimen is on the left.

The smooth fatigue pre-crack extends approximately 0.1 inch to

the right and distinctly ends. The crack extends via slow fatigue

crack growth until about 0.125 inch from the free surface, when

final unstable fracture occurs.

Specimen 19 was loaded for 16,548 additional cycles past

aluminum plate failure prior to laminate failure. Laminate

failure was dominated by a disbond extending from the crack

plane to the upper laminate taper tip. The disbond occurred

between the resin and fabric of the interface lamina with resin

still bonded to the aluminum plate. Strain data indicates initial

laminate delamination initiates at approximately 36,000 cycles

and continues to grow until ultimate failure at 50,871 cycles.

Inter-lamina failure initiated between the layer of 7500 fabric

and ±45 plies and resulted in the delamination of the 7500 fabric

from the adhesive bondline. The adhesive remained adhered to

the aluminum surface, as shown in Figure 12. Figure 13 shows

the specimen 19 aluminum fracture surface for the strain gauge

0/1 ligament. The center notch of the specimen is on the left.

The smooth fatigue pre-crack extends approximately 0.25 inch

to the right and distinctly ends. The crack extends via slow

fatigue crack growth until about 0.25 inch from the free surface,

when final fracture occurs. Application of the composite patch

repair increased the fatigue life of the 11 inch wide cracked

plate by 7.5 and 6.5 times for the as-delivered and sensitized

aluminum, respectively.

Figure 10: Specimen 10 post fatigue disbond extents marked with the dashed red line (L) and arrows mark extents of laminate bridging on the strain gauge 2/3

ligament (R).

Figure 11: Specimen 10 post fatigue strain gauge 0/1 aluminum fracture surface.

Figure 12: Specimen 19 composite failure surface showing resin and fibers adhered to the aluminum

surface.

Figure 13: Specimen 19 (patched, sensitized aluminum) fracture surface for strain gauge 0/1

ligament.

14.5 ksi Far-Field Stress Level The 14.5 ksi far-field stress level testing of specimens 14,

12, and 18 was performed at a load cycle rate of 1 Hz from 4 to

40 kips, or a far-field stress range of 1.5 to 14.5 ksi. Strain

gauge data was offset by the first load segments zero load strain

data. Each subsequent set of strain data was offset using the

same zero data.

The un-patched, as-received aluminum specimen 14

failed at a cycle count of 1,047. Ligament failure occurred first

on the 0/1strain gauge side followed by failure on the strain

gauge 2/3 side. The center notch of the specimen is on the left of

the photograph and extends to a distinct end approximately 0.25

inch to the right. Failure surfaces appear to primarily be slow

speed fatigue crack growth until failure of the final 2.0 inches of

ligament. The failure surface on the strain gauge 0/1 ligament is

shown Figure 14.

Slow Speed Fatigue FractureFatigue Fracture Unstable Fracture


Figure 14: Specimen 14 fracture surface on the strain gauge 0/1 ligament.

Aluminum failure occurred at cycle counts of 16,390 and

12,517 for composite patch repaired specimens 12 (as-received

aluminum) and 18 (sensitized aluminum), respectively. The

composite patch repairs failed shortly after aluminum plate

failure. Specimen 12 loading continued for an additional 268

cycles prior to laminate failure, which consisted of failure

between the 0/90 ply 3 and the ±45 ply 4. Laminate failure is

shown in Figure 15 with the extent of delamination marked by

the dashed red line. Localized failure of the 7500 ply occurred

adjacent to the crack; however failure propagated along the ply

3 and 4 interface as seen in Figure 16.

Specimen 18 was loaded for 364 additional cycles past

aluminum plate failure prior to laminate failure. Strain data

indicates laminate failure initiated with the final aluminum

failure and progressed along the adhesive to 7500 ply bondline

until the final failure at cycle 12,881. The patch failure is shown

in Figure 17; the adhesive remained adhered to the aluminum

surface. Figure 18 shows the specimen 18 aluminum fracture

surface for the strain gauge 0/1 ligament. The center notch of the

specimen is on the right, with the fatigue pre-crack extending to

the left approximately 0.25 inch to the left. Slow fatigue crack

growth occurs until about the final 0.25 inch of the ligament,

when final fracture occurs. Application of the composite patch

repair increased the fatigue life of the 11 inch wide cracked

plate by 15.7 and 12.0 times for the as-delivered and sensitized

aluminum, respectively.

Figure 15: Specimen 12 post fatigue failures. Extent of delamination marked in the red line and the black

dashed line is the initial crack.

Figure 16: Specimen 12 laminate failure.

Figure 17: Specimen 18 laminate failure surface between the resin and the 7500 fabric with resin still

bonded to the aluminum surface.

Figure 18: Specimen 18 aluminum fracture surface of the strain gauge 0/1 ligament.

TEST RESULTS SUMMARY Fatigue testing results show that the composite patch

repair system increases the fatigue resistance of the aluminum

baseline specimens. Firm conclusions with statistical

significance cannot be made using this data; however general

trends and the overall performance of the composite patch repair

system and the as-received to sensitized aluminum performance

may be noted. In general the sensitized aluminum plate had

lower fatigue cycles to failure, though more specimens are

required to confirm the trend. Also notable was the failure

surface of the sensitized aluminum shown in Figure 13, typically

the final unstable fracture surface appeared shiny and smooth

for the oven sensitized aluminum.

Test details and results for specimens with an initial crack

length of 5 inches are listed in Table 3. Stress level as a function

of failure cycle for the un-patched aluminum, composite patch

repaired aluminum, and the oven sensitized composite patch

aluminum for the three stress levels are shown in Figure 19. The

Disbonded

Laminate

Lamina

Bonded to

Aluminum

Resin

Bonded to

Aluminum


data points for each plate condition are fit with a power law to

highlight the general trend of the data and to show the difference

between the plate condition groups. Each data point represents

one specimen.

Application of a composite patch repair to a cracked

aluminum plate subjected to a cyclic tensile load increased the

cycle count to failure for the three stress levels tested. The

benefit of the composite patch repair system increased with the

level of applied stress. When subjected to a far-field tensile

stress level of 5 ksi the number of load cycles increased by 4 to

5 times for the sensitized and as-delivered plate condition.

Similarly, for the 10.9 ksi stress level the cycle counts increased

6.5 to 7.5 times. When the stress level was increased to 14.5 ksi

the cycle count increased between 12 to 15.7 times the baseline

cracked plates.

Composite patch disbond, when it occurred, was limited

to a small area adjacent to the crack plane and extended less

than 0.5 inches beyond the crack tip. For the 5.0 and 10.9 ksi

stress levels the composite still carried load immediately

following failure of the aluminum plate. When composite failure

occurred, the adhesive remained bonded to the surface of the

aluminum plate.

Under a far-field tensile stress level of 5 ksi the number

of load cycles to aluminum failure increased by 4 to 5 times for

the sensitized and as-delivered plate condition. Similarly, for the

10.9 ksi stress level the cycles to aluminum failure increased 6.5

to 7.5 times. When the stress level was increased to 14.5 ksi the

cycles to aluminum failure increased between 12 to 15.7 times

the baseline un-patched plates

Table 3: Fatigue Test Results for Specimens 2a0 = 5 inch.

Figure 19: Far-field stress level versus cycles to failure for 2a0 = 5 inch fatigue testing.

FINITE ELEMENT ANALYSIS Analytical predictions were made for a far-field stress of

5 ksi and correlated with test data for specimens 16, 11, and 17;

the aluminum, composite patch repaired aluminum, and the

composite patch repaired oven-sensitized aluminum specimens

respectively.

The HYSEFDA toolkit is based on extended finite

element (XFEM) theory and implemented with phantom paired

element approach. HYSEFDA predicts material fracture by

allowing crack insertions that are independent of the finite

element mesh with user defined elements with XFEM

capabilities. Delamination between the aluminum and composite

patch was predicted using 2-noded user defined elements.

Originally designed to model fiber-metal laminate (FML)

structures, HYSEFDA is capable of capturing the interaction

between metal and composite and its effect on the growth of

both metallic crack and interfacial delamination under fatigue

loading.

Test specimen geometry was modeled using linear solid

elements and symmetry along the loading axis. The cross

section mesh of the 3D finite element model is shown in Figure

20, where the green elements represent the aluminum and the

gray elements are the composite patch. The composite patch was

modeled with 8-noded continuum shell elements with the

composite layup defined within the section definition. The

properties for individual lamina were obtained from either direct

testing of individual plies or by scaling manufacturer provided

properties in classical lamination theory (CLT) to match

laminate tensile testing data. Table 4 summarizes the stiffness

properties used for each ply, with standard aluminum properties

assumed; 10 Msi modulus and a Poisson’s ratio of 0.3. Only

Paris’ Law parameters are required for crack and delamination

growth prediction; parameters used are listed in Table 5. The

initial crack was explicitly modeled as an initial condition plane

through enriched elements, as shown in Figure 21, had an initial

half crack length of 2.5 inches.

Spec ID

Peak Load

(lbs)

Far-Field

Stress (psi)

Cycles to

Aluminum Failure

Fatigue Cycle

Increase

16 14000 5091 54628 ----

11 14000 5091 284300 5.2

17 14000 5091 234800 4.3

15 30000 10909 5305 ----

10 30000 10909 39890 7.5

19 30000 10909 34323 6.5

14 40000 14545 1047 ----

12 40000 14545 16390 15.718 40000 14545 12517 12.0

Fatigue Specimens (2a = 5 inches )

0

2000

4000

6000

8000

10000

12000

14000

16000

1 10 100 1,000 10,000 100,000 1,000,000

Far-

Fie

ld S

tre

ss (

psi

)

Cycles

CCT Specimen Fatigue Cycles to Aluminum Failure5 Inch Crack Length

Aluminum Only

Patched

Sensitized Patched


Figure 20: Solid FEM developed for HYSEFDA toolkit.

Table 4: Lamina mechanical properties.

Ply Lamina

Architecture

E

(msi)

G12

(msi) v12

Hexcel

7500 Plain weave 2.83 0.56 0.3

Vectorply -45/45 1.51 0.56 0.3

Vectorply 0/90 1.51 0.56 0.3

Hexcel

7781 Plain weave 2.83 0.56 0.3

Table 5: Paris’ Law parameters.

C (ksi-mm) n

Aluminum Fracture 8.0e-9 3.0

Interfacial Delamination 5.1e5 7.5

Figure 21: Location of explicitly modeled crack extending from the plane of symmetry.

Figure 22 shows the aluminum crack tip (top) and

composite patch delamination (bottom) when the predicted

crack propagation was 1.5 inches, for a half crack length of 4

inches. Because user defined elements are used to predict the

fracture, the location of the crack tip and extent of delamination

must be visualized using the maximum principle strain values in

the aluminum elements. The crack propagated in a mostly

straight line perpendicular to the load axis and the predicted

delamination damage extended outward 0.3 inches from the

crack plane along the crack path.

Figure 22: Propagated crack and interfacial delamination at half crack length of 4.0 inches.

Simulation results show a similar life extension benefit

for the composite patched specimen, which is highlighted by the

crack length as a function of cycle count responses (a-N) shown

in Figure 23. Table 6 compares the cycles-to-failure results from

numerical simulation to those from experiments. It can be seen

that numerical simulation predictions correlate well with the

limited test data available. The predicted cycles to failure for the

aluminum CCT specimen were within 2% of the test results,

while the prediction for the composite patched CCT specimens

was 12 to 27% lower than reported for testing. The cycles-to-

failure predicted for the composite patched specimens were

conservative due to the assumed symmetry of the finite element

model (FEM). Asymmetrical failure was observed during

testing; one ligament failed followed by the second. This pattern

was possibly due to imperfect fixtures or slight fixture

alignment, neither of which were measured and therefore not

represented by the FEM. When comparing the initiation of

failure, characterized by the first significant non-linearity of the

CMOD versus cycle count response the simulation correlates

well with the test data. Table 7 and Table 8 compare the crack

mouth opening displacement (CMOD) as a function of cycles

for both cases with and without the patch. Good agreement is

obtained between numerical and experimental results. Again,

very good agreement was achieved for the aluminum CCT

specimen.


Table 6: Cycles to failure comparison between experiment and simulation.

Cases

Far field

Stress

(ksi)

Experimental

result

Numerical

result

Without

Patch 5.0

54,628

(Specimen 16) 55,800

With

Patch 5.0

234,800

(Specimen 17) 205,000

284,300

(Specimen 11)

Figure 23: Crack length versus cycle for CCT with and without a composite patch.

Figure 24: CMOD comparison between experiment and simulation without a composite patch.

Figure 25: CMOD comparison between experiment and simulation with a composite patch.

To demonstrate the accuracy in determining the crack

growth driving force for a patched specimen, a verification

study has been performed on the numerical model by comparing

the HYSEFDA prediction of the stress intensity factor at the

initial crack length with an analytical solution and a stationary

crack model using virtual crack closing technique (VCCT) in

ABAQUS. Stress intensity at the crack tip for cases without a

patch and with a patch is summarized in Table 7 and Table 8,

respectively. For the case without patch, an analytical solution

exists(Anderson, 1995). For the case with a single sided

composite patch, the stress intensity and crack opening

displacement are not uniform through the plate thickness due to

bending. A FEM cross section along the load axis with

magnified displacements is shown in Figure 26. In the figure,

the composite patch is represented by the tan layer and the

aluminum plate is the light green layer. Magnified

displacements show the variation in the crack opening

magnitude from the patched surface to the free surface, which

translates into varying crack tip stress intensity through the

thickness. In Table 8 the stress intensity is reported for the crack

tip at the patched surface and the free surface, opposite of the

composite patch. It can be seen that for all cases the HYSEFDA

predictions agree with the analytical and the VCCT solutions.

Based on this verification study, we can assume that the

discrepancy shown in Figure 25 can be attributed to the assumed

geometric symmetry and the uncertainty of fatigue properties

selected for this simulation.

Table 7: Comparison of K at the initial crack length for an aluminum plate.

Model K (ksi·in0.5

)

Analytical 16.1

HYSEFDA 15.94

Abaqus VCCT 15.98

Table 8: Comparison of K at the initial crack length for a composite patch repaired aluminum plate.

Model Patch Surface

K (ksi·in0.5

)

Free Surface K

(ksi·in0.5

)

HYSEFDA 0.43 1.26


ABAQUS

VCCT 0.43 1.30

Figure 26: Load axis cross section showing crack opening magnified 20x. (Tan-Composite, Light Green-

Aluminum)

CONCLUSIONS Testing of the composite patch repaired CCT specimens

demonstrated the proof of concept and the performance of the

repair at three far-field stress states and an initial crack length of

5 inches and the initial correlation of analytical predictions for

the 5 ksi far-field stress level. Repair system performance was

measured and analyzed under tension-tension fatigue loading

with as-delivered and oven sensitized aluminum plate to

complete the following three objectives:

1. Measure the tension-tension fatigue cycles to aluminum

failure for both aluminum and composite patch repaired

specimens.

2. Observe and document the behavior of the composite

patch repair laminate during aluminum crack

propagation.

3. Compare composite patch repair performance on

sensitized and as-delivered aluminum plate.

4. Correlated the analytical and test response of CCT

specimens to tensile fatigue loading at the 5 ksi far-field

stress level.

Application of a composite patch repair to a cracked

aluminum plate subjected to a cyclic tensile load increased the

cycle count to failure for the three stress levels tested. The

benefit of the composite patch repair system increased with the

level of applied stress. Performance increases observed are of

the same magnitude reported by (Schubbe, 1999) for smaller

specimens at a far-field stress level of 17.4 ksi using uni-

directional boron/epoxy patches.

Damage propagation with the quasi-isotropic layup and

rubber toughened epoxy was limited to a small area adjacent to

the crack plane for the 5 and 10.9 ksi stress levels. The limited

damage area and the residual resin on the aluminum are a partial

measure of the durability of the repair system.

Tension-tension fatigue testing of the composite patch

repaired aluminum plates beyond aluminum plate failure

demonstrated the ability of the composite patch repair to support

tensile, in-plane stress levels up to 10.9 ksi, which is above the

ABS aluminum fatigue design limit of 7.5 ksi. Though more

research and development is required for application to marine

structures, demonstrating the load capacity makes utilizing

composite patches for not only repairs but also structural

reinforcement an option for designers.

HYSEFDA analytical response predictions correlated

well with the measured test response for the 5 ksi applied far-

field stress. Though correlating well with the low far-field stress

level the HYSEFDA toolkit will be evaluated further on the full

geometry to predict asymmetry and then used to predict the

fatigue response of the higher stress levels where more plastic

and ductile material response was observed.

ACKNOWLEDGEMENTS Authors would like to acknowledge and thank Dr. Paul

Hess from ONR for funding and guidance on the project and Dr.

Jim Lua of GEM for his guidance with the analytical

predictions. Special thanks to John Noland, Tim Dapp, and John

Kim of NSWCCD who performed specimen manufacturing,

instrumentation, and editing. Without their help this work would

not have been possible.

WORKS CITED

ABS. (2009). ABS Guide for Building and Classing

Naval Vessels. ABS Americas.

Anderson, T. (1995). Fracture Mechanics

Fundamentals and Applications, Second Edition.

CRC Press LLC.

Baker, A. R. (2002). Advances in the Bonded

Composite Repair of Metallic Aircraft Structure (Vol.

1). Amsterdam, The Netherlands: Elsevier.

Bone, J. K. (2003). Testing of COmposite Patches for

Ship Plating Fracture Repair. Masters Thesis, Naval

Architecture and Marine Engineering, University of

Michigan.

Grabovac, I. (2003). Bonded Composite Solution to

Ship Reinforcement. Composites Part A: Applied

Science and Manufacturing , vol. 34, 847.

Hosseini-Toudeshki H., M. B. (2007). Numerical and

Experimental Fatigue Crack Growth Analysis in

Mode-I for Repaired Aluminum Panels Using

Composite Material. Composites: Part A , 38, 1141-

1148.

Noland, J. H. (2013). Initiatives in Bonded Ship

Structural Repairs. American Society of Naval

Engineers Fleet Maintenance and Modernization

Symposium. San Diego, CA.

O'Shaugnessy, T. (Nov 1985). Green Letter No. 143

ALCOA 5000-Series Alloys Suitable for Welded

Structural Applications. Pennsylvania 15069:

Aluminum Company of America Application

Engineering Division ALCOA Laboratories Center.

Sabelkin, V. M. (2006). Fatigue Crack Growth

Analysis of Stiffened Cracked Panel Repaired with


Bonded COmposite Patch. Engeineering Fracture

Mechanics , 73, 1553-1567.

Schubbe, J. M. (1999). Investigation of a Cracked

Thick Aluminum Panel Repaired with a Bonded

Composite Patch. Engineering Fracture Mechanics ,

63, 305-323.

Weitzenbock, J. M. (2012). A Cold Repair Method

for FPSOs. Offshore Technology Conference - 23543.

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