quantification and correlation of pit parameters to … · agard proceeding on "fatigue in the...

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QUANTIFICATION AND CORRELATION OF PIT PARAMETERSTO "SMALL" FATIGUE CRACKS

Dr. V. Chandrasekaran1

Ms. A.M.H. Taylor2

Mr. Young-In Yoon3

Prof. D.W. Hoeppner4

The primary objective of this study was to correlate corrosion"damage" using pit depth as a controlling factor with the "small"fatigue crack growth rates of 2024-T3 aluminum alloy specimensincluding a specimen made from a dismantled JSTAR fuselagepanel. In all fatigue experiments, the cycles to first detectable cracklength were recorded. A video recording system was used toobserve the specimen surface, to record real time video as well asto measure cracks during fatigue experiments. The initialdetectable crack length recorded using this system was in the range0.15 to 0.21 mm. From the plots of ‘a’ vs. N for all of thespecimens, it can be observed that the smaller the range of pitdepth, the longer the fatigue cycles to form the first detectablecrack. Moreover, in general, the crack growth rates for the prior-corroded specimens and the JSTAR specimen were greater than theuncorroded specimen. At low ∆K values (for example @ 0.1MPa√m and 0.25 MPa√m), the crack growth rates of the prior-corroded specimens in the "small" region were considerably fasterthan the uncorroded specimen.

Finally, fractographic analysis revealed some interestingresults. The origins of the cracks were as expected for the mostpart. On the uncorroded specimens the origins were at the cornerof the rivet hole where the stress concentration was highest, due tono other damage on the surface. The prior corroded specimenscracked at the rivet hole also at the points of large corrosion pits.The most important result was the crack origin of the JSTARspecimen, which was inside the rivet hole. This shows the extentof the damage inside the rivet holes that may go unseen but thatmay result in failure.

1 Research Assistant Professor, Mechanical Engineering Department, University of Utah, Salt Lake City,Utah 84112.2 Research Engineer, Mechanical Engineering Department, University of Utah, Salt Lake City, Utah 84112.3 Doctoral Student, Mechanical Engineering Department, University of Utah, Salt Lake City, Utah 84112.4 Professor and Director, Quality and Integrity Design Engineering Center (QIDEC), MechanicalEngineering Department, University of Utah, Salt Lake City, Utah 84112.

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INTRODUCTION

As aircraft age, the processes of corrosion and corrosion fatigue are expected todegrade the structural integrity of aircraft structures. Corrosion acting conjointly withfatigue can have major effects on materials in structures of aircraft. Corrosion can creatediscontinuities (pits, cracks, etc) that act as origins of fatigue cracks with significantreductions in life at all stress levels. As Wallace and Hoeppner mentioned in theirAGARD report on "Aircraft Corrosion: Causes and Case Histories", in the initial stages,corrosion is in the form of filiform or pitting in the interior and exterior of fuselage skins[1]. Moreover, in the aging aircraft fleet, pitting corrosion has recently been viewed as asafety concern as it "may accelerate the onset of widespread fatigue damage" [2]. A recentAGARD proceeding on "Fatigue in the Presence of Corrosion" addressed the significanceof corrosion on the structural integrity of aircraft structures [3]. As well, the failureanalysis of a RAAF F/A-18 trailing edge flap concluded that corrosion pits in the 7050-T73 flap outboard hinge lug nucleated fatigue cracks that resulted in the final fracture ofthe subject flap [4]. Moreover, in our study, when a specimen extracted from adismantled JSTAR fuselage panel was analyzed using the confocal microscope, numerouspits were observed around a rivet hole as shown in Figure 1.

In crack propagation, corrosion effects are well known to produce acceleratedfatigue crack propagation. Pitting corrosion fatigue models have been proposed in thepast to incorporate pitting parameters to predict the nucleation of a Mode I crack frompits and also to facilitate crack propagation analysis [5-9]. Also, a procedure forestimating fatigue lives of corrosion-pitting-induced airframe components was proposed[10].

Although major efforts were expended to understand the crack propagationbehavior of materials, a few studies have focused on the crack nucleation stage in theoverall fatigue process [5,11,12]. In 1928, McAdam first suggested that corrosion inducedpits might act as stress concentrators from which cracks could form [13]. A large numberof chemical or electrochemical factors such as potential, passive film, pH, andcomposition of environment are found to affect the pitting corrosion fatigue process.Mechanical factors such as stress range, frequency, stress ratio (R), and load waveformand metallurgical factors such as material composition, microstructure, heat treatment,and orientation can influence pitting corrosion fatigue process. Nucleation of cracks fromcorrosion pits were observed by many researchers including the works of Hoeppner[5,11,12], Goto [14] in heat-treated carbon steel, and Muller [15] in several steels. Aswell, in NaCl environment, lowering of the fatigue life due to the generation of pits incarbon steel [16] and 7075-T6 aluminum alloy [17] was observed under corrosion fatigueconditions.

Experimental study was performed to characterize quantitatively the effect ofcorrosion damage on the "small" fatigue crack growth rates of 2024-T3 aluminum alloyspecimens. In this study, a "small" fatigue crack is defined as a crack that is <1 mm inlength. Corrosion pits were artificially produced on 2024-T3 aluminum alloy specimens.Experimental details are discussed in the following section.

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EXPERIMENTAL DETAILSTEST SPECIMENS

Fatigue specimens were machined from 1.52 mm (0.06 inch) thick 2024-T3aluminum alloy sheet material. The configuration of the specimen is shown in Figure 2.Fatigue specimens were also machined from a discarded 2024-T3 aluminum alloy JSTARfuselage panel. The fatigue specimens (except the JSTAR test specimens) were polishedusing silicon carbide abrasive grit 200, 320, 400, and 600 size papers with distilled wateras a lubricant, followed by a rough polishing with 1 micron diamond paste and finalpolishing with 0.3 micron diamond paste. Then, the specimens were prepared forartificial corrosion to produce pits on the surface. The specimens were masked withparafin wax except around the machined hole at the center section of the specimen. Themachined hole was also filled with wax to avoid corrosion inside the hole. As well, thecurved edges at the center section of the specimen were masked to prevent corrosionalong the edges. The test specimens made from the JSTAR fuselage panel already had arivet hole. The prior corrosion technique used to produce pits on the specimen surface isdiscussed in the next section.

PRIOR CORROSION METHOD

A solution comprising 5% of NaCl was prepared in 500 ml of deionized water.The specimen was immersed in the solution bath with a carbon rod as the cathode. A DCvoltage of about 1-2V was applied across the surface of the specimen to accelerate theformation of corrosion pits on the exposed area of the specimen. The corrosion processwas found to begin within a few seconds after the voltage was applied. The depth of pitswas varied by changing the exposure time of the specimens in solution. The time ofexposure of specimens in solution was varied from 4 to 8 minutes. By this technique a pitdepth in the range from 5 to 15 µm was obtained on the corroded area of the specimens.After attaining the desired corrosion damage, the specimens were ultrasonically cleanedwith acetone and stored in a desiccation chamber until the fatigue testing was started.Figure 3 shows the nature of pits produced on one of the prior-corroded specimens testedin this study. Prior to testing, the confocal microscope was used to quantify the depth ofpits that were artificially produced on 2024-T3 aluminum alloy specimens as described inour earlier report [18]. For the JSTAR specimen, the range of pit depth was quantifiedusing the confocal microscope in the "as is" condition. Figure 4 shows the surfacecorrosion damage around the rivet hole of the JSTAR specimen tested in this study.Details of fatigue test set up are discussed below.

FATIGUE TEST SET UP

Fatigue experiments were performed using an MTS servo controlled electro-hydraulic horizontal fatigue machine. A recording system was used to observe thespecimen surface, to record real time video as well as to measure cracks during fatigueexperiments. A schematic illustration of the system is shown in Figure 5. This system

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uses a microscope with a zoom lens fitted to an Hitachi KP-112 video camera. Thespecimen surface could be magnified between 29X and 372X on a panasonic 13"monitor. The microscope has two degrees of freedom. It was placed parallel to thespecimen surface. The specimen surface was monitored in-situ during fatigueexperiments and recorded in real time. The crack length measurement was possible usinga JV-6000 video micrometer. As the specimen could only be mounted horizontally on thefatigue machine, only the upper surface of the specimen was monitored during fatigueexperiments. The primary objective of this study was to correlate the quantified pit depthwith the "small" fatigue crack growth rates of plain, prior-corroded or prior-pitted, andthe JSTAR 2024-T3 aluminum alloy specimens. The fatigue experimental results arediscussed in the following section.

FATIGUE TEST RESULTS

A total of six fatigue experiments were performed on 2024-T3 aluminum alloyspecimens: one test on a plain (uncorroded) specimen, four tests on prior-corrodedspecimens and one test on a JSTAR specimen. The test parameters and the results aregiven in Table 1. As shown in Table 1, the greater the range of pit depth, the lower thefatigue life of 2024-T3 aluminum alloy specimens.

Table 1 - Fatigue Test Parameters and Test Results

(All experiments were conducted at room temperature (laboratory air)

Specimen ID Material Test Parameter Range ofPit Depth(µm)

FatigueCycles atFracture

Plain oruncorroded

2024-T3 σmax = 138 MPa (20ksi, Net stress); R =0.1, f = 10 Hz

Notapplicable

291,200

Prior-corroded#1or prior-pitted #1


5-10 122,100



5-10 106,400



5-15 29,000



5-15 92,000

JSTAR 2024-T3 σmax = 138 MPa (20ksi, Net stress); R =0.1, f = 10 Hz

2-6 188,700

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Moreover, in all fatigue experiments, the cycles to first detectable crack lengthwere recorded. The initial detectable crack length recorded using the system describedbefore was in the range 0.15 to 0.21 mm. From the results shown in Table 2, it can beobserved that in general, the smaller the range of pit depth, the longer the fatigue cycles toform the first detectable crack. In addition, it was observed that the number of fatiguecycles to form the first detectable crack for the uncorroded specimen was much higherwhen compared to the prior-corroded as well as the JSTAR specimen. Therefore, thenumber of fatigue cycles to a detectable crack size at a given stress level is accelerated byprior-corrosion in the 2024-T3 aluminum alloy specimens tested in this study. This aloneindicates that prior-corrosion indeed has an effect in the fatigue crack formation process.

Table 2 - Cycles to the first detectable through crack for uncorroded and prior-corroded2024-T3 aluminum alloy specimens

Specimen ID Range ofPit Depth(µm)

Length offirstdetectablecrack (mm)

Fatiguecycles tofirstdetectablecrack

Finalcracklengthmeasured(mm)**

Fatigue cyclesto fracture

Plain oruncorroded

Notapplicable

0.201 70,200 2.849 291,200


5-10 0.157 20,000 2.653 122,100


5-10 0.149 20,100 2.679 106,400


5-15 0.189 5000* 2.398 29,000


5-15 0.154 20,000 2.417 92,000

JSTAR 2-6 0.216 35,000 2.507 188,700*Tested at σmax = 207 MPa (σnet = 30 ksi);. All other tests were conducted at σmax = 138MPa (σnet = 20 ksi);. R = 0.1, f = 10 Hz, Room temperature (laboratory air).** That is when unstable fracture occurred.

As shown in Table 2, the length of the final crack that was measured before thefracture became unstable, varied in relation to the quantified range of pit depth. The finalcrack length measured on the uncorroded specimen was considerably longer whencompared to that of the prior-corroded and the JSTAR specimens.

To reduce the raw data (‘a’ vs. N) to da/dN vs. ∆K, stress intensity solution forthis particular specimen geometry (refer Figure 2) was needed. It was provided by APESIncorporated. The stress intensity values are for a through crack coming out of the holeand they are given in Figure 6. Even though crack(s) formed and propagated from bothsides of the hole during fatigue testing, the lead crack was always found to grow only

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from one side of the hole until that side fractured. Therefore, the crack length datagenerated in this study are from one side of the hole.

DISCUSSION OF EXPERIMENTAL RESULTS

The results from the fatigue experiments were as expected. From the results of 'a'vs. N for all of the specimens, it can be seen that the smaller the range of pit depth, thelonger the fatigue life. The uncorroded specimen lasted the longest of them all. Acomparison of ‘a’ vs. N data in both the "small" and long crack region for all thespecimens is shown in Figure 7. As shown, the crack lengths in the "small" crack region(<1mm) are close to linear. As the crack transitioned into the long crack region the cracklengths grew exponentially. Moreover, as shown in Figure 7, the ‘a’ vs. N curves for thespecimens show a gradual shift to the left indicating a decrease in the overall fatigue lifewith the increase in the quantified range of pit depth. Also, when a prior-corrodedspecimen (range of pit depth 5-15 µm) was tested at σmax of 207 MPa (σnet = 30 ksi), itresulted in a greater reduction in the fatigue life when compared to the other specimenwith the same range of pit depth that was tested at σmax of 138 MPa (σnet = 20 ksi).

The crack growth rates for the prior-corroded specimens and the JSTAR specimenwere faster than the uncorroded specimen. The uncorroded specimen most closelyfollowed the typical shape of a da/dN vs. ∆K curve. However, the da/dN vs. ∆K results ofthe prior-corroded and the JSTAR specimens exhibited considerable scatter as expected.This might be due to the fact that it was sometimes difficult to follow the crack(s) on theprior-corroded specimens because of all the surface damage caused by the pits. Figure 8shows a comparison of da/dN vs. ∆K results of all of the specimens tested in thisprogram. As shown, at low ∆K values (for example @ 0.1 MPa√m and 0.25 MPa√m), thecrack growth rates of the prior-corroded specimens in the "small" region wereconsiderably faster than the uncorroded specimen. Moreover, as shown in Figure 8, at a∆K value of 0.1 MPa√m, the "small" crack growth rates of the JSTAR specimen wasfound to be faster than the uncorroded, as well as, the prior-corroded specimens with therange of pit depth of 5-10 µm. However, at the same ∆K value, the "small" crack growthrates of the JSTAR specimen was slower than the prior-corroded specimen with 5-15 µmpit depth. Subsequently, the AFGROW program was used to perform a preliminaryanalysis as discussed in the next section.

AFGROW ANALYSIS OF EXPERIMENTAL RESULTS

The AFGROW program was used to perform a preliminary analysis of theexperimentally generated data using the Harter-T method with the "user defined" throughcrack model. To use the "user defined" through crack model, the AFGROW programneeded the beta values for each specimen. Also, it needed an "initial surface crack length"(C) for analysis. As mentioned before, in all fatigue tests, the ‘first detectable throughcrack’ for each specimen was measured and it was used as the input for ‘C’. Theestimated number of fatigue cycles to fracture by AFGROW for all of the specimens isgiven in Table 3. In general, the AFGROW program over-predicted the fatigue life of

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prior-corroded 2024-T3 aluminum alloy specimens. However, the fatigue life estimationby the AFGROW for the uncorroded and the JSTAR specimens was conservative.

From the results shown in Table 3, it appears that the estimation of fatiguelife by the AFGROW program is largely dependent on the specification of the value forthe "initial surface crack length (C)". However, the presence of corrosion damageincluding the pits on the specimen surface may alter the crack propagation behavior inboth the "small" and long crack region under fatigue loading conditions. This may be thereason why the AFGROW program over-estimated the fatigue life of the prior-corrodedspecimens.

Table 3 - Comparison of AFGROW results with fatigue experimental results

Specimen ID Range ofPit Depth(µm)

Length of the firstdetectable crack(mm)/Initial cracklength 'C'

Fatigue cycles tofracture(AFGROW result)

Fatigue cyclesto fracture (Testresult)

Plain oruncorroded

Notapplicable

0.201 183,700 291,200


5-10 0.157 147,000 122,100


5-10 0.149 149,000 106,400


5-15 0.154 148,500 92,000

JSTAR 2-6 0.216 144,100 188,700

SEM ANALYSIS OF FRACTURE SURFACES

A fractographic examination was performed on the failed specimens to determineany differences of the fracture surfaces between the uncorroded specimens, the specimenswith prior corrosion damage, and the specimens taken from the JSTAR fuselage.Examination of the fracture surfaces was conducted visually with the naked eye and witha Hitachi S-2300 Scanning Electron Microscope (SEM). Many unique features werefound between the three types of specimens on the fracture surfaces and on the exteriorsurface where the corrosion damage was or was not present.

SEM ANALYSIS OF UNCORRODED SPECIMENS

Although the uncorroded specimens were not subjected to any corrosion damage,it can be seen in Figures 9-11 how 2024-T3 aluminum is naturally full of pits ordiscontinuities, on the surface and throughout the material. These pits may act as naturalstress concentrations in the material.

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The uncorroded specimens exhibited areas with fatigue striations, as seen inFigure 9. Figure 10 shows the crack origin on one side of the uncorroded specimen. Thisfigure also shows river patterns pointing back to the origin. River patterns are typicalsigns of fatigue crack growth, as well as the transgranular facets visible in this figure.

Finally, Figure 10 shows the transition from fatigue crack growth to fast fracture,visible by the dimples on the right side of the figure. Figure 11 displays the fracturesurface of the uncorroded specimen, as well as the exterior surface and inside the rivethole. It may be seen that these surfaces are free from any damage, other than typicalmachining marks. Fractography of the uncorroded specimens did not produce anysurprises in the failure analysis. Typical fatigue crack growth and fast fracturecharacteristics were present.

SEM ANALYSIS OF PRIOR CORRODED SPECIMENS

Figures 12-14 show the extent of the exterior surface damage on the priorcorroded specimens. Figure 12 shows the variety of pits on the surface of specimen #1;deep and shallow, small and large. The crack origin of prior corroded specimen #2 isshown in Figure 13. It is clear from this figure and further examination of the fracturesurface, that the crack originated from a large pit at the corner of the rivet hole. Asecondary crack between two pits was captured in Figure 14. This crack shows how pitswill produce cracks that may link up to cause failure under fatigue loading even withoutthe presence of a rivet hole.

Figure 15 shows a crack origin in a corrosion pit at the corner of the rivet hole.This figure shows how deeply the corrosion damage penetrated into the surface of thematerial. The prior corroded specimens exhibited fracture characteristics similar to thoseof the uncorroded specimens, such as fatigue striations, transgranular facets, and dimples.Figure 16 displays the clear transition from fatigue crack growth to fast fracture in a semi-circular pattern. Facets and dimples are shown in each region, respectively.

SEM ANALYSIS OF JSTAR FUSELAGE SPECIMENS

The exterior surfaces of the specimens from the JSTAR aircraft fuselage weredifferent from the prior corroded specimens because they were naturally corroded anddamaged in service. Figures 17-19 show this. The bright spots on the pictures are spotson the material that are electrically charging in the SEM. This is often caused by residualcorrosion product, as may be seen on the prior corroded specimens in Figures 12 and 13.

Figures 17-19 demonstrate how damaged the exterior surfaces of the specimenswere with pits, scratches, deep grooves, and surface cracks (Figure 19). Figure 20displays the transition from the fatigue crack propagation region to the fast fracture regionon one side of JSTAR specimen. River patterns point back to the crack origin at the top,left of the figure. A profile of the surface damage on the specimen is also visible at thetop of the micrograph. A crack origin inside the rivet hole is shown in Figure 21 on theother side of JSTAR specimen. The damage inside the rivet hole that could cause a crackto form there may be seen. Figure 22 shows a close-up of the crack origin. The charginginside the rivet hole on the right may indicate corrosion product.

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Figures 23 and 24 show the fracture surface of JSTAR specimen. Fatiguestriations are evident. The natural pits in aluminum alloy 2024-T3, mentioned in theuncorroded specimen section, are also present.

Overall, the crack fronts of all of the specimens were similar. On the side wherethe crack propagated to a longer length, and which eventually broke first, the crack frontswere straight. On the side with the smaller crack length, the crack fronts were semi-circular. This side failed in overload, as was seen from the ductile dimples present on thefracture surface. Figure 20 shows the semi-circular crack front and the transition from theregion of fatigue crack growth and fast fracture.

All of the specimens exhibited fracture surface characteristics typical of fatiguecrack growth. Fatigue striations, transgranular facets, and river patterns were visible onall of the specimens. The origins of the cracks were as expected for the most part. On theuncorroded specimens the origins were at the corner of the rivet hole where the stressconcentration was highest, due to no other damage on the surface. The prior corrodedspecimens cracked at the rivet hole also at the points of large corrosion pits. The mostimportant result was the crack origin of JSTAR specimen, which was inside the rivethole. This shows the extent of the damage inside the rivet holes that may go unseen butthat may result in failure.

CONCLUSIONS

From this experimental study the following conclusions can be made:

! The smaller the range of pit depth, the longer the fatigue cycles to form the firstdetectable crack length. In addition, it was observed that the number of fatigue cyclesto first detectable crack for the uncorroded specimen was much higher whencompared to the prior-corroded, as well as, the JSTAR specimen. As well, the lengthof the final crack that was measured before the fracture became unstable varied inrelation to the quantified range of pit depth. The final crack length of the plainspecimen was considerably longer when compared to that of the prior-corroded andthe JSTAR specimens.

! The crack growth rates for the prior-corroded specimens and the JSTAR specimenwere greater than the uncorroded specimen. In addition, the greater the range of pitdepth, the lower the fatigue life of the specimens.

! At low ∆K values (for example @ 0.1 MPa√m and 0.25 MPa√m), the crack growthrates of the prior-corroded specimens in the "small" region were considerably fasterthan the uncorroded specimen.

! In general, the analysis of the experimentally generated data using the AFGROWprogram resulted in the over-prediction of the fatigue life of prior-corroded 2024-T3aluminum alloy specimens. However, the fatigue life estimation by the AFGROW forthe uncorroded and the JSTAR specimens was conservative.

! The fractography analysis revealed some interesting results. On the side where thecrack propagated to a longer length, and which eventually broke first, the crack frontswere straight. On the side with the smaller crack length, the crack fronts were semi-

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circular. This side failed in overload, as was seen from the ductile dimples present onthe fracture surface.

! The origins of the cracks were as expected for the most part. On the uncorrodedspecimens the origins were at the corner of the rivet hole where the stressconcentration was highest, due to no other damage on the surface. The prior corrodedspecimens cracked at the rivet hole also at the points of large corrosion pits.

! The most important result was the crack origin of JSTAR specimen, which was insidethe rivet hole. This shows the extent of the damage inside the rivet holes that may gounseen but that may result in failure.

! From the experimentally generated results, it is thereby concluded that prior-corrosionindeed affects the fatigue crack growth rates of 2024-T3 aluminum alloy specimens inthe "small" crack region, as well as, altering the cracking mechanism(s) as observedfrom fractographic analysis.

ACKNOWLEDGEMENT

The research reported herein was performed as a part of the research program"Modeling corrosion growth on aircraft structure" for NCI Information Systems, Fairborn,OH, under contract NCI USAF 9138-003 with Mr. Garth Cooke as the program manager.The support of NCI Information Systems in conducting this research is gratefullyacknowledged. Also, we greatly appreciate the help of Dr. Scott Prost-Domasky of APES,Inc. for providing the stress intensity solution and also his assistance in performingAFGROW analysis.

REFERENCES

1. Wallace, W. and Hoeppner, D.W. "Aircraft Corrosion: Causes and Case Histories",AGARD Corrosion Handbook, Vol. 1, AGARD-AG-278-Vol. 1, 1985.

2. Lincoln, J.W., "Corrosion and Fatigue: Safety Issue or Economic Issue," in Fatigue inthe Presence of Corrosion, Papers presented at the Workshop of the RTO AppliedVehicle Technology (AVT) Panel (organised by the former AGARD Structures andMaterials Panel) held in Corfu, Greece, 7-9 October 1998, RTO-MP-18, Publisher:Research and Technology Organization, BP 25, 7 Rue Ancelle, F-92201 Neuilly-Sur-Seine Cedex, France, 2-1 to 2-5, March 1999.

3. Fatigue in the Presence of Corrosion, Papers presented at the Workshop of the RTOApplied Vehicle Technology (AVT) Panel (organised by the former AGARDStructures and Materials Panel) held in Corfu, Greece, 7-9 October 1998, RTO-MP-18, Publisher: Research and Technology Organization, BP 25, 7 Rue Ancelle, F-92201 Neuilly-Sur-Seine Cedex, France, RTO-MP-18, March 1999.

4. Barter, S., Sharp, P.K., and Clark, G., "The Failure of An F/A-18 Trailing Edge FlapHinge," Engineering Failure Analysis, Vol. 1, No. 4, 255-266, 1994.

5. Hoeppner, D.W., “Model for Prediction of Fatigue Lives Based Upon a PittingCorrosion Fatigue Process,” Fatigue Mechanisms, Proceedings of an ASTM-NBS-

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NSF Symposium, J.T. Fong, Ed., ASTM STP 675, American Society for Testing andMaterials, 1979 pp. 841-870.

6. Lindley, T. C., McIntyre, P., and Trant, P. J., “Fatigue Crack Initiation at CorrosionPits,” Metals Technology, Vol. 9, 1982, pp. 135-142.

7. Kawai, S. and Kasai, K., “Considerations of Allowable Stress of Corrosion Fatigue(Focused on the Influence of Pitting),” Fatigue Fracture of Engineering MaterialsStructure, Vol. 8, No. 2, 1985, pp. 115-127.

8. Kondo, Y., “Prediction of Fatigue Crack Initiation Life Based on Pit Growth,”Corrosion Science, Vol. 45, No. 1, 1989, pp. 7-11.

9. Wei, R.P., Li, C., Harlow, D.G., and Flournoy, T.H., "Probability Modelling ofCorrosion Fatigue Crack Growth and Pitting Corrosion," in Fatigue in New andAgeing Aircraft, ICAF 97, Proceedings of the 19th Symposium of the InternationalCommittee on Aeronautical Fatigue 18-20 June 1997, Edinburgh, Scotland, Vol I,197-214, 1997.

10. Perez, R., "Corrosion/Fatigue Metrics," in Fatigue in New and Ageing Aircraft, ICAF97, Proceedings of the 19th Symposium of the International Committee onAeronautical Fatigue 18-20 June 1997, Edinburgh, Scotland, Vol I, 215-230, 1997.

11. Hoeppner, D.W., Mann, and Weekes, “Fracture Mechanics Based Modelling of

Corrosion Fatigue Process,” in Corrosion Fatigue: Proceedings of the 52nd meeting ofthe AGARD Structural and Materials Panel held in Turkey, 5-10 April, 1981.

12. Hoeppner, D.W., Corrosion Fatigue Considerations in Materials Selections andEngineering Design”, Corrosion Fatigue: Chemistry, Mechanics, and Microstructure,NACE, 1972, pp. 3-11.

13. McAdam, D.J., and Gell, G.W., “Pitting and Its Effect on the Fatigue Limit of SteelsCorroded Under Various Conditions”, Journal of the Proceedings of the AmericanSociety for Testing Materials, Vol. 41, 1928, pp. 696-732.

14. Goto, M., and Nisitani, H., “Crack Initiation and Propagation Behavior of a Heattreated Carbon Steel in Corrosion Fatigue”, Fatigue Fracture Engineering MaterialStructure, Vol. 15, No. 4, 1992, pp. 353-363.

15. Muller, M., “Theoretical Considerations on Corrosion Fatigue Crack Initiation”,Metallurgical Transactions, Vol. 13A, 1982, pp. 649-655.

16. Mehdizadeh, P., et al., “Corrosion Fatigue Performance of a Carbon Steel in BrineContaining Air, H2S, CO2”, Corrosion, Vol. 22, 1966, pp. 325-335.

17. Corsetti, L.V., and Duquette, D.J., “The Effect of Mean Stress and Environment onCorrosion Fatigue Behavior of 7075-T6 Aluminum”, Metallurgical Transactions,Vol. 5, 1974, pp. 1087-1093.

18. Hoeppner, D.W., and Chandrasekaran, V. "Correlation of Pit Depth to Fatigue Life of2024-T3 Aluminum Alloy Specimens - An Experimental Study", FASIDE report toNCI Informations Systems Inc., Fairborn, OH, November, 1998.

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Figure 1: Digitized confocal microscope image showing numerous pits around a rivethole of a specimen extracted from a JSTAR fuselage panel (Range of pit depth 5-20 µm).

Figure 2: Fatigue specimen configuration (All dimensions in mm).

25.40

R = 12.70 +0.07 -0.00Ø9.53+0.07

-0.00

TWO HOLES

76.20

50.80

18.42+0.07-0.00

1.52+0.07-0.00

C

C

Ø5.00+0.10-0.10

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Figure 3: SEM micrograph showing artificially produced pits on 2024-T3 aluminum alloyspecimen tested in this study (X80).

Figure 4: Photograph showing surface corrosion damage on the JSTAR specimen testedin this study.

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Figure 5: Schematic illustration of the recording system and test setup

Video micrometer

Character generator

VCR Monitor

Video printer

Fiber optic illuminator

Microscope

Ring illumination light

Specimen

Mounting system

Video camera

PinGrips

(Cyclic Load)Load cell

end Actuator end

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Figure 6: The stress intensity values for the specimen geometry used in this study.

Figure 7: Summary chart of 'a' vs. N data in both "small" (<1 mm) and "long" crackregion for all the 2024-T3 aluminum alloy specimens.

'a' vs. N data ("small" and long regime) for 2024-T3 aluminum alloy specimens (Plain as well as prior-pitted specimens)

0

0.5

1

1.5

2

2.5

3

0.0E+00 5.0E+04 1.0E+05 1.5E+05 2.0E+05 2.5E+05 3.0E+05 3.5E+05

Fatigue Cycles (N)

Cra

ck L

eng

th (

mea

sure

d in

mm

)

Material: 2024-T3 aluminum alloyMax. Net stress = 138 and 207 MPa (20 ksi and 30 ksi)R = 0.1Frequency = 10 Hz.

1. Plain specimen , 138 MPa (20 ksi)2. JSTAR Fuselage panel specimen, 138 MPa (20 ksi) (2-6 micron pit depth)3. Prior-pitted specimen, 138 MPa (20 ksi) (5-10 micron pit depth)4. Prior-pitted specimen , 138 MPa (20 ksi) (5-15 micron pit depth)5. Prior-pitted specimen , 207 MPa (30 ksi) (5-15 micron pit depth)

4 32

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Cycles to first detectable through crack

Stress intensity factor (KI) values for one crack (through) from the hole for the specimen configuration used in this study

0

200

400

600

800

1000

1200

0.00 0.50 1.00 1.50 2.00 2.50 3.00

Crack Length 'a' in mm

K in

MP

a sq

rt. m

m

16

Figure 8: Summary chart of 'da/dN' vs. ∆K data in both "small" and long crack region forall the 2024-T3 aluminum alloy specimens.

Figure 9: Fatigue striations on uncorroded specimen (X800).

Fatigue Crack Growth Rate Data for 2024-T3 Aluminum Alloy Specimens

1.00E-10

1.00E-09

1.00E-08

1.00E-07

1.00E-06

1.00E-05

0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0

Delta K (MPa sqrt. m)

Lo

g d

a/d

N (

m/c

ycle

)

Series1

Series2Series3Series4Series5

Material: 2024-T3 aluminum alloyMax. Net stress = 138 MPa (20 ksi)R = 0.1Frequency = 10 Hz.Room temperature (Lab. air)

Series 1:Plain or uncorroded specimen

Series 2 & 3:Prior-pitted specimen (5-10 micron pit depth)

Series 4:Prior-pitted specimen (5-15 micron pit depth)

Series 5:JSTAR Fuselage panel specimen (2-6 micron pit depth)

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Figure 10: Crack origin and fracture features of baseline specimen (X200).

Figure 11: Fracture surface and exterior surface and rivet hole of uncorroded specimen(X200).

Crack origin

Inside rivet hole

Exteriorsurface

Fracture surface

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Figure 12: Corrosion pits on the surface of prior-corroded specimen #1 (X200).

Figure 13: Crack origin at corrosion pit on prior-corroded specimen #2 (X200).

Inside rivethole

Exteriorsurface

Crack origin at corrosion pit

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Figure 14: A crack between corrosion pits on the surface of prior-corroded specimen #2(X1.0k).

Figure 15: Crack origin in a deep corrosion pit on prior corroded specimen #4 (X200).

Crack originat a pit

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Figure 16: Transition from fatigue crack growth to fast fracture on prior corrodedspecimen #4 (X150).

Figure 17: Surface damage on JSTAR specimen (X100).

Direction of fracture mode transition

21

Figure 18: Surface damage on JSTAR specimen (X200).

Figure 19: Surface damage on JSTAR specimen with surface cracks visible (X400).

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Figure 20: Transition from FCG to fast fracture on JSTAR specimen (X60).

Figure 21: Crack origin inside rivet hole on JSTAR specimen (X150).

Crackorigin

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Figure 22: Close up of crack origin (see Figure 21 also) on JSTAR specimen (X800).

Figure 23: Fatigue striations on JSTAR specimen fracture surface (X1.0k).

Crack origin

Corrosionproduct insiderivet hole asindicated byelectricalcharging in theSEM.

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Figure 24: Fatigue striations on JSTAR specimen fracture surface (X1.0k).

quantification and correlation of pit parameters to … · agard proceeding on "fatigue in the...

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