Paper No. Year-2014 Hart 1

Fatigue Performance and Analysis of Composite Patch Repaired Cracked Aluminum Plates

Author Name(s): D. C. Hart1 (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 & McGeorge, 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, Karr, & Waas, 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, Mall, & Avram, 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 (Schubbe & Mall, 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, Mohammedi, Sadeghi, & Daghyani, 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 watertight integrity to the compromised compartment and mitigating or stopping crack growth. The initial composite patch repair design was based on Classical Lamination Theory (CLT), Rose’s Model closed form solution (Baker, Rose, & Jones, 2002), and simple Finite

Paper No. Year - 2014 Hart 2

Element Analyses (FEA) reducing the stress intensity at the aluminum crack tip by more than 25% (Noland, Hart, Loup, Udinski, & Sielski, 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 was 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) (Fang, Stuebner, & Lua, 2013).

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 were 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 were 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”

Paper No. Year - 2014 Hart 3

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

Aluminum As Delivered 10.9 0.1 15

Aluminum w/Patch As Delivered 10.9 0.1 10

Aluminum w/Patch Oven Sensitized 10.9 0.1 19

Aluminum As Delivered 14.5 0.1 14

Aluminum w/Patch As Delivered 14.5 0.1 12

Aluminum w/Patch Oven Sensitized 14.5 0.1 18

Total 9

Paper No. Year - 2014 Hart 4

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

Paper No. Year - 2014 Hart 5

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

CMOD (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

CMOD (in)

Cycle

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

Paper No. Year - 2014 Hart 6

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.

Paper No. Year - 2014 Hart 7

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. The reason for this requires further investigation.

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

Paper No. Year - 2014 Hart 8

aluminum for the three stress levels are shown in Figure 19. The 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 15.7

18 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‐Field Stress (psi)

Cycles

CCT Specimen Fatigue Cycles to Aluminum Failure5 Inch Crack Length

Aluminum Only

Patched

Sensitized Patched

Paper No. Year - 2014 Hart 9

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 (Link, 1988).

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.

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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: Simulated 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

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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 & Mall, 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.

Future efforts are required to advance naval use of composite patches beyond low stress planar applications. These should include testing and analysis such that analytical methods and design guidance can be developed to address non-structural repair to structural alteration applications.

More CCT testing of composite patched shipboard sensitized plate will produce a complete data set with as-delivered, oven-sensitized, and shipboard sensitized plate; demonstrating performance for unchanged, heat affected, and both heat and corrosion affected plate conditions. In addition to CCT testing the load and structural configuration should increase in complexity to cover cases such as pressure and bending.

Research and development of analytical methods to capture failure modes associated with the marine structures will produce the design guidance required to extend current composite patch repair from low stress applications to structural repairs and class wide alterations.

Current needs for composite patches include both aluminum and steel structures. Applications include plate wastage, plate cracking, structural damage, and structural reinforcement. Plate wastage affects decks, hulls, and tanks and often requires costly crop out and replace repair methods. Cracking due to sensitization, fatigue and overload can be reduced with the use of composite patches, or bonded repair, and significantly reduce the cost of maintenance and increase service life.

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.

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

Vessels. ABS Americas.

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