Fatigue Crack Growth in 7050T7451 Aluminium Alloy Thick Section ...

Fatigue Crack Growth in 7050T7451 AluminiumAlloy Thick Section Plate with a Glass Bead PeenedSurface Simulating Some Regions of the F/A-18StructureExecutive SummaryAVD research into material factors likely to affect F/A-18 fatigue life has highlightedthe critical role played by surface condition in determining the service fatigue life ofaircraft structure. This report presents the results of a fatigue coupon test program.The primary purpose was to obtain results from coupons with a glass bead peenedsurface condition typical of some regions of the critical structure of the F/A-18 aircraft.Another surface condition that could possibly initiate critical cracks was compared tothese results.These specimens were loaded with a representative wing root-bending spectrumderived from the F/A-18 FT488/2 centre barrel bulkhead fatigue test. Several peakstress levels were used. The coupons were representative of the material and geometrytypical of a structural detail that has been found to be fatigue-critical in the 7050T7451high strength aluminium alloy wing carry through bulkheads. The spectrum used hadadditional marker loads added to aid quantitative fractography. These marker loadsare briefly discussed.Following the tests, quantitative fractography was used to produce crack growthcurves for each of the fatigue specimens. This allowed a detailed interpretation of thecrack growth to be made, including a measure of the severity of the flaws from whichthe fatigue cracks initiated. In addition to an examination of the effect of the flaws, itwas found that the peening produced a retarding effect on the early part of the crackgrowth. The extent of this effect was examined by comparison to previously examinedetched specimens. These comparisons are reported.

DSTO-TR-14771. IntroductionDSTO research into material factors likely to be of importance to the F/A-18 fatigue life havehighlighted the critical role played by surface condition in determining the service fatigue life ofaircraft structure. Previous coupon test programs have addressed the differences betweenpolished, ‘as-machined’ and peened surfaces in terms of life when loaded by several differentwing root bending moment spectra; Molent et al , 2000; Sharp & Clark, 2001. This report examinesthe application of glass bead peening to aluminium alloy 7050T7451 thick section plate typical ofareas in the F/A-18 which have been treated to extend their fatigue life. Work similar to this haspreviously been carried out and reported in Sharp & Clark, 2001 for some wing root bendingspectra. The specimens reported on in this paper were loaded with more representative wing rootbending spectrum derived from the F/A-18 FT55 centre fuselage fatigue test being carried out inCanada, Simpson, 2002. Whereas the previous testing compared lives at a single peak spectrumstress, this examination reports on several peak stresses in order to establish the trend in the LifeImprovement Factor (LIF). In addition, a more detailed examination of the crack growthmorphology including detailed quantitative fractography on the largest cracks, at these differentpeak spectrum stresses is presented. The use of several peak stress levels produced life dataapplicable in the estimation of the LIFs for the most critical structure in the F/A-18 aircraft. TheFT55 spectrum used had additional marker loads added to aid quantitative fractography.This project was developed out of a desire to improve the understanding of fatigue crackingencountered in service and test components from the F/A-18 aircraft. To simulate the peenedsurface of the AL7050T7451 material in the aircraft, the surfaces of the test coupons were preparedto simulate those of 7050 components fitted to the F/A-18.2.1 Introduction2. BackgroundPeening treatments are widely used in mechanical and aeronautical engineering to improve thefatigue lives of components. The process is normally associated with harder metals such as steels,nickel and titanium alloys. Although this is generally the case, simply configured aluminium alloycomponents have for many years been peened, for example aluminium alloy propeller blade hubs.The application of this method to complex aluminium alloy components such as those found inthe F/A-18 is however a relatively recent occurrence. This has resulted in reports of extensivevariation in the fatigue life results for these high strength aluminium alloy peened componentsand, in some cases (Clayton & Clark, 1988), a decrease in fatigue life has been observed. Suchvariability naturally raises concern that the peening process developed for use with hard materialssuch as steels might not always be suitable for peening high-strength aluminium alloys which aremuch softer.An extensive array of literature dating back to early in the twentieth century examining the effectsof peening on steels and other hard alloys exists. An excellent review of this literature may befound in the book produced by the Metal Improvement Company, Inc of New Jersey, USAentitled “Shot Peening Applications” eighth edition, 2001. In a general sense most of this isrelevant although the softness of the aluminium alloys compared to these materials results insome special problems which need to be addressed by an improved understanding of the wayfatigue cracking propagates in a peened aluminium alloy.1

DSTO-TR-1477To build the required level of understanding, DSTO has over the years undertaken a number ofresearch programs investigating the effects of steel shot, glass and ceramic bead peening on 7050aluminium alloy - an alloy which is used extensively in RAAF F/A-18 aircraft. The overall aim ofthis previous work was to establish a LIF for the peening undertaken on various parts of the F/A-18. This previous work also attempted to provide a means of measuring peening quality byassessing the surface condition after peening, and thereby the reliability of the fatigue lifeimprovement, Sharp & Clark, 2001. While the flaws that initiated cracking in the peening wereunderstood to be an important factor in the effectiveness of the peening, no attempt was made toquantify the way the fatigue cracks grew from them.2.2 Peening and its relevance to the F/A-18Peening attempts to deform material surrounding the impact point of a projectile impacting thetarget surface, against the resistance of the material surrounding the impact. In so doing, acomplex sub-surface residual stress distribution is generated in which, generally, the material justbelow and to the sides of the impact are in elastic compression as shown in Figure 1, produced bythe enlarged surface layer. To balance this compression zone an elastic tension zone surroundsthe compression zone. The transition to this tension zone can be rapid. For the case were thesurface has been covered by impacts the zones may be considered to be in layers receding fromthe surface so that the surface is generally in compression followed by a layer of tension which, onmoving deeper decays to zero as shown in Figure 2. These deformation induced residual stressesare triaxial with a balanced biaxial state and as such will affect the apparent loading of anydiscontinuity, such as a fatigue crack, that is growing through the layers. The compressiveresidual stresses keep the crack closed when a far field tensile stress is applied to a component, toa level related to the magnitude of the compressive residual stresses, reducing the amount of cracktip opening. This reduction in crack opening, and therefore crack tip stress intensity range (∆K)retards the cracks growth. The retardation caused by these compressive stresses still operateswhen the crack tip has grown into the tensile stress field at deeper depths, where the crack tip isheld open by the tensile residual stresses since the crack behind the tip (wake) is still clampedclosed by the compressive residual stresses. Therefore the effect of the residual stresses producedby peening is always to retard the crack growth by raising the far field stress required to open thecrack from its far field unloaded state.ImpactingprojectileFigure 1Surface of metalRegion of material held incompression by surroundingmaterialThe process of deforming the surface as the result of the impact of a projectile produces residualstrain which is more commonly referred to as residual compressive stress. From “Shot PeeningApplications”, 2001.2

DSTO-TR-1477Figure 2Schematic of residual stress distribution below a peened surface. With traditional bead peeningtechniques this compressive layer generally extends 0.2-0.4mm below the surface (from Sharp &Clark, 2001).In reality peening does not produce a consistent surface condition since the resultant residualstresses are the product of many individual impacts by the peening media. These impacts are notdistributed evenly over a surface, so the result is a variation in the final residual stress field. Anexample of a peened 7050T7451 surface as imaged by backscattered electron channellingdiffraction pattern mapping methods in the scanning electron microscope (SEM) is shown inFigure 3. Here the variation in colour shows the variation in the intensity of the residual strains.Surface20 micronsFigure 3Example of strain variation in a peened section of 7050T7451 aluminium alloy as shown bybackscattered electron channelling diffraction pattern mapping. The strain appears to be highestdirectly below the centre of a peening dent (arrow). The lines indicate the sub-grain boundaries.(Note that the image does not show the full extent of the peening, just the variation in strain)Along with the above local variations, the residual stress field is reliant on the constraint that ismaintained by the material surrounding the deformed material. This constraint disappears at afree surface, and can be reduced at external radii (re-entrant radii usually have excellentconstraint). This leads to the most favourable region for crack propagation being at externalcorners of peened components. Since the surface is unconstrained the compressive residual stressusually drops very close to the surface and, because the surface is not smooth, may containregions of residual tension. This allows surface flaws to initiate fatigue cracks more favourably3

DSTO-TR-1477than internal flaws. An example 1 of the measured residual stresses for a glass bead peened7050T7451 specimen is shown in Figure 4.500-50-100-150-200-2500 0.2 0.4 0.6 0.8 1 1.2Depth below the surface mm.Figure 4Measured residual stresses through the surface layers of a 7050T7451 specimen peened withglass beads, corrected to show the stresses at a depth of 1mm as zero, Sharp & Clark, 2001.2.3 Peening in the F/A-18Many critical areas of the F/A-18 structure are peened. Three different forms of peening havebeen used: glass bead, ceramic bead and steel shot 2 . Generally OEM peening was carried out withsteel or glass beads whereas areas that are now peened are typically peened, or re-peened in somecases with ceramic beads. DSTO tests on the difference in the lives produced by these threemethods on the critical 7050T7451 material have indicated that steel shot results in significantsurface damage which, under some conditions will result in lives that are shorter than theunpeened lives. It has been found that although the fatigue crack growth rate is retarded in thecompressive residual stress layer produced by the peening, the damage to the surface from thesteel shot peening (laps folds and embedded shot) produce crack like defects that can be so largeas to shorten the total crack growth life to below the unpeened life even with the retardationincluded, Clayton and Clark, 1988, Clark and Clayton, 1991 and Sharp and Clark, 2001. On theother hand glass bead peening produces a shallower residual stress layer along with the reductionin damage, producing a combination that usually leads to an extension to the fatigue life, since thedamage produced by this process is usually considerably shallower than the extent of the depth towhich crack retardation occurs. This trade off between surface damage and crack retardationdepth can still become marginal due to large flaws being produced by impact with fractured1 The residual stress distribution shown in Figure 4 is not necessarily applicable to all glass bead peened7050T7451 surfaces. The depth to which the residual stresses for glass bead peened 7050T7451 reach aremore typically about 0.25mm on a flat surface and a little less at corners. This of course depends on whatis used as the surface: the top of the peened areas or the base of the peening dents. The variation in thestrain noted in Figure 3 and the difficulties in the measurement of residual stresses also adds to theuncertainty in residual stress plots.2 Flapper wheel peening is included in the McDonnell Douglas process specification for peening: PS14023.1, although it is not recommended for fillet areas and was recommended ‘for peening previouslypeened areas and for assemblies requiring rework’. Glass beads were originally recommended foraluminium alloy peening in this specification.4

DSTO-TR-1477beads, fragments of which may become embedded in the surface, excessive dwell times due tomanual application of the peening process resulting in deep folds and laps, and incorrect peeningof corners – folding over sharp corners by peening towards the edge, then burring this materialaround the edge preventing the surface beneath the burr from being peened.To reduce the risk of bead fracture while maintaining the controllability of glass beads, ceramicbeads were introduced. These are denser and stronger than glass while being lighter than steelshot. Tests comparing ceramic bead peened specimens to glass bead peened specimens found thaton average they gave a similar life – the difference was that the glass bead peened specimens hada larger scatter in fatigue life as the result of the damage caused by fractured beads (Sharp et. al.,1994). Given that glass bead peening if correctly used, generally gives reasonable lifeimprovements, and that several of the critical areas of the F/A-18 aircraft structure have beenglass bead peened, this method was adopted to prepare the surfaces of the coupons tested andreported on here.3.1 Introduction3. Fatigue testingA series of fatigue tests designed to represent typical RAAF service loading were carried out. Thespecimens were of a type used previously to compare spectra developed during the lead-up to thetesting of full-scale F/A-18 structure both in Canada and Australia (DSTO) under the IFOSTP,Simpson et. al., 2002. These previous tests also included tests carried out with etched surfaces,surfaces similar to those of F/A-18 fracture critical 7050 components, using the same spectrum asused here, Barter, 2003. Prior to the etched surface specimen testing, testing was either carried outusing different spectra or unrepresentative surface conditions. In the case of the peened specimens(Sharp and Clark, 2001) testing was mostly aimed at disclosing some of the critical factors inproducing a reliably peened surface rather than establishing the LIF for varying stress levelsunder a representative spectrum.3.2 Specimens geometryThe specimens were of the simple ‘dog bone’ type with a continuous 127mm radius on either sideresulting in a 15mm wide test section. The thickness of the specimens was 6.35mm giving a crosssectionalarea of 95.25mm 2 at the specimen centre. A schematic of the specimens is shown inFigure 5. The K t of 1.037 for the radius was calculated by FE analysis. This specimen was originallydesigned to mimic a critical low K t detail of an F/A-18 Y488 centre-section bulkhead, for which alot of data has already been generated using this specimen or similar specimens of a closelyrelated configuration. Although the Y488 bulkhead detail in question has a 6” (152.4mm) radius,the K t was not significantly different (Molent, Ogden and Pell, 2000) and the 127mm radiusallowed a shorter specimen that used less material. Experience with fatigue cracking in thesespecimens indicated that critical cracks could occur up to about 15mm on either side of thespecimen centre. Using, for simplicity a constant width of 10mm over this region an estimate ofthe approximate surface area likely to generate a critical crack in each specimen was about850mm 2 . Therefore for each stress level with five specimens tested about 4270mm 2 of peenedsurface was tested. Over this area the specimen varies in width from 15mm at the centre to about16mm at 15mm from the centre. This will result in about a 5.6% drop in the net section appliedstress at 10mm from the centre. Countering this, the rounding of the corners prior to peeningresulted in a measured drop in cross-sectional area of between 1 and 3.5%. More on the criticalregions for peened surfaces will be presented later.5

DSTO-TR-1477127mm Radius (5’)230mm65mm15mm6.35mmFigure 5A drawing of the specimens used in this testing programThe specimens were machined from a 5.5 inch thick 7050-T7451 aluminium alloy plate with theDSTO, AVD designation of ‘KS’ (produced by Ravenswood Aluminium Corporation,Ravenswood, West Virginia 3 ). The long axis of the specimen aligned with the rolling direction andthe width of the specimen in the rolling plane as shown, in Figure 6. This gave the coupons a grainorientation that was more resistant to fatigue crack growth than would have been the case hadthey been cut with the long axis across the plate width. The cracking in these specimens will be inthe Short/Transverse plane. The structural detail that was the subject of the previous testinggenerates cracking consistent with ‘across the plate’ specimens – cracking in theShort/Longitudinal plane, therefore the specimens tested here are not strictly representative of themould line flange of the F/A-18 bulkheads.Rolling directionT (Transverse)S (Short Transverse)L (Logitudinal)Figure 6A schematic showing the orientation of the specimens as they were cut from the plate3.3 Specimen Material PropertiesThe KS plate was certified by Ravenswood to conform to the requirements of the specificationAMS 4050E which is very similar to the McDonnell Douglas Process specification 1420. Thechemical composition specification requirements are shown in Table 1 along with the results ofanalysis of the KS plate. Table 2 shows the mechanical properties and conductivity requirementsof 7050T7451. The supplied mechanical properties and conductivity measurements are shown inTable 3 along with the measured conductivity and hardness values.3 now – Century Aluminium Company6

DSTO-TR-1477Table 1 Specification AMS 4050E – Chemical composition %.Element Min Max Analysis of KSplateZinc 5.7 6.7 6.42Copper 2 2.6 2.20Magnesium 1.9 2.6 1.99Zirconium 0.08 0.15 0.08Iron -- 0.15 0.06Silicon -- 0.12 0.03Manganese -- 0.1

DSTO-TR-1477At the surface large grains were not disguisable. These larger grains tended to surround the finer,equiaxed grains (10-20µm) that made up the bulk of the microstructure.3.4 Surface PreparationTo prepare the specimens for glass bead peening, the specimen edges were rounded to a 1 to 2mmradius and the surfaces were polished to a 600# finish. Peening was carried out in a repeatablemanner using a mechanical scanning frame. The corners were first peened followed by each of thefaces. Peening was to an intensity of Almen 6-8A (SAE J442 standard), after 200% coverage. Thebeads were all virgin - none were reused. This method is similar to that used to repair regions ofthe F/A-18 aircraft and could be expected to produce similar life results. The grip regions of thespecimens were also peened to prevent failure in the grip area.3.5 Loading and Marker loadsThe loading applied to the specimens was a modified IARPO3a spectrum as applied to FT488/2,Athiniotis, et, al, 2003. It had been derived from the FT55 wing root bending moment spectrumapplied to the Canadian F/A-18 centre fuselage test article. The IARPO3a spectrum used in thistesting is equivalent to 324.92 Simulated Flight Hours (SFH) and consisted of 13,475 turningpoints, 13,480 including marker loads (see below). The original spectrum was developed fromflight data collected from the wing root strain gauges from service aircraft.Several peak load levels were selected and the spectrum was factored accordingly. The peak loadswere selected to represent some of the peak loads expected in the centre section bulkheads of theF/A-18 in Australian service. The highest load level was selected to be greater than the highestknown loading experienced by several details in the F/A-18. This was carried out to cover anypossibility that these load levels would be reached in service either through ‘spikes’ or increases inloading due to repair of critical areas, where peening is most likely to be used. The test matrix isshown in Table 4. Since there is a small Kt associated with the shape of the test section of thespecimen the stress at the root of the radii will have been a little larger than the net section stress.This along with the net section stress difference associated with failure of the specimen away fromthe centre (common in these peened specimens), and the slight difference in the cross sectionalareas due to the rounding of the corners of the specimens (calculated to average about 2%) hasbeen ignored in the presentation of the data. Further analysis of the effects of these stressvariations may lead to further comment in a following report. Loading was applied at 10Hz with acomputer controlled servo hydraulic fatigue test machine.Table 4 Test matrixPeak Spectrum Stress Surface Condition No. of specimensMPa (Ksi)tested360 (52.2) Glass bead peened 5390 (56.6) “ 5420 (60.9) “ 5450 (65.3) “ 5Since quantitative fractography (QF) was to be carried out on the fractures produced during thetesting it was decided to add marker-loads to the spectrum. After several tests of different markerload additions, five compressive marker loads were found to be the best compromise. These wereinserted into the spectrum just before one of the spectrum peak loads, between the four precedingtensile loads. A graphical representation of these marker loads is shown in Figure 7.8

DSTO-TR-14773.6.1 Crack morphologyAll the main fatigue cracks produced during this phase of the testing were examined by QF andthe crack growth rates measured (reported in Section 3.7.2). As part of this examination severalgeneral observations about the appearance of the crack surfaces, and the secondary cracks, can bemade. The most obvious difference between the etched specimens examined in Barter, 2003, andthe peened specimens was the number of larger cracks. While numerous cracks had initiated inthe etched specimens, and many of these cracks were not far behind the growth of the crack/s thatcaused failure, this was not the case in the peened specimens. The peened specimens had very fewsignificant cracks, although a large number of much smaller cracks did exist. In general only a fewsignificant cracks intersected the fracture surface, and in some cases only a single significant crack(Figure 10A). This led to some very jagged fracture surfaces (Figure 10B) due to the fast fracturerunning along the grain orientation rather than linking together significant cracks during failure.This can be compared with the large number of cracks usually found on the fracture plane of oneof the etched specimens, Figure 11. The distribution of cracks over the surface of the specimenswas reasonably even although the size of the cracking compared to the main crack/s was usuallyvery small. Some of these small secondary cracks could usually be seen in the stretch zones(surface regions that yielded during failure of the specimen) either side of the failure plane.Examples are shown in Figure 12 .ABFigure 10 The fracture in one of the peened specimens ‘A’ (390MPa loading) showing the single crack fromwhich failure occurred. Note that since the cracking was across the rolling direction the fracturewas heavily distorted along the rolling direction as shown in ‘B’.11

DSTO-TR-1477Figure 11 The fracture surface from one of the etched specimens. Note the large number of fatigue cracks(darker grey areas) that have grown around the entire circumference of the fracture surface.ABCFigure 12 The stretch zone surface adjacent to the fractures of several peened specimens showing the manysecondary cracks. ‘A’ has some of these cracks marked with arrows and ‘B’ shows peeningdamage from which a larger crack started marked with an arrow.D12

DSTO-TR-1477The largest cracks were found growing from all positions (sides, edges and corners up to about10mm either side of the centre of the specimen) from the specimen surface, although thepredominant position was at or near the external radii in the specimen edge (5”radius) as shownin Figure 13. Here an example of a main crack with two main regions of initiation, both of whichare just within the radius of the corner which was a favoured site for initiation for many of themain cracks. Although the corners were favoured as sites for initiation, these corners had beenrounded with a 1 to 2mm radius to prevent them from being overly significant as initiation sitesand to mimic the type of peening rework carried out on F/A-18 structure.Figure 13 Typical region of crack initiation near a corner. This was the most common position for crackinitiation in these peened specimens. Note that the initiation sites are not in the middle of thecurved section but closer to the transition region.The number of cracks that could be observed intersecting the fracture appeared to increased withthe stress level and specimen life (at a particular stress level).The crack growth around the origins of the cracks, particularly the cracks which occurred in thespecimens loaded at the lower stress levels revealed considerable distortion of the normal crackfront shape. Some examples are shown in Figure 14. Typically the cracks initially grew with alonger surface length than depth compared to cracks in the etched specimens (Barter, 2003), thenrapidly deepened resulting in a bulging out of the centre of the crack front. This is evident inFigure 14B. (The banding in this Figure shows the repeat in the block loading). The shape was tosome extent distorted in this manner due to the deepest point of the crack moving into thematerial that is less affected by the residual stresses produced by the peening, while the remainderof the crack fount is still affected by the residual stress. Odd shaped crack fronts can also occurduring cracking from cold worked fastener holes, Clark, 1991.13

DSTO-TR-1477A (KSIF152)B (KSIF107)Figure 14 Two views of the regions about initiations of the fatigue cracking in the peened specimens. ‘B’shows the approximate position of one of the marker bands – red dashed line, indicating the oddshaped crack front while the crack was small. The arrows point out some of the block repeats.Crack initiation was in all cases found to be associated with flaws in the surface, as was the casewith the etched specimens. Most usually the origins consisted of closely spaced multipleinitiations along the edges of laps produced during peening. These were usually the mostprevalent type of flaw in the glass bead peened surface and were found, to a grater or lesserdegree, to depend on the quality of the peening, as has been noted previously by Sharp et al, 2001.An example of such an origin site is shown in Figure 15A. Those laps that are most likely toinitiate cracking are usually at the base of peening dents. The second view in Figure 15B shows adeep cut produced by a fractured glass bead. This also produced multiple origins around its edge.Although multiple origins initially occurred the crack quickly formed a single crack front, whichprogressed more rapidly into the material than it did at its flanks.14A (KSIF193)B (KSIF142)Figure 15 Two examples of the types of flaws that initiate fatigue cracks in peened 7050T7451. ‘A’ showsthe closely spaced multiple initiations associated with a broad lap produced during peening(approximately outlined). ‘B’ shows a deep cut and associated damage produced by a fracturedglass bead impact (approximately outlined). This also produced multiple origins around its edge.

DSTO-TR-1477Sections through one of the specimens surfaces showing examples of cracks growing from a lapand a sharp glass bead cut are presented in Figure 16.Figure 16 Sections through the surface of one of the specimens, showing examples of fatigue cracks growingfrom a lap and a sharp glass bead cut3.7 Quantitative fractography3.7.1 IntroductionQuantitative fractography (QF) was carried out on the largest crack in each specimen. Themethod of calculating the crack depth during measurement of crack growth is set out in Barter,2003. These analyses were made possible by the existence of a fairly distinct repeating patternaided by the marker loads as discussed earlier. The repeating markings were found to be veryconsistent with the type and spacing expected for fatigue growth per spectrum under the testconditions. Usually the repeating pattern from the repeat of the spectrum was fairly easily foundparticularly around the originating flaws, although its distinctiveness did vary from area to areaon the fracture surface. By tracking the most distinct areas and some of the less distinct areas ofthe repeating patterns and combining the data from these regions, a relatively complete picture ofthe crack growth could be assembled assuming that each repeat of the pattern was equivalent to asingle load spectrum. This is supported by previous examinations of fatigue fracture surfaces in7050 aluminium alloy, which have undergone similar fatigue testing, Barter, 1990a, Barter, 1991,Bishop and Clark, 1991, Barter, 1998, Barter and Price, 2000.For the peened specimens, the crack shape could be very distorted from the semi-elliptical shapenormally found in the absence of residual stresses, as noted in the previous section. In all thesecases the QF was carried out on the region as close as possible to the centre and deepest part of thecrack to avoid artefacts in the crack growth curves.3.7.2 Quantitative fractography measurementsFigure 17A&B to Figure 20A&B present the crack growth curves produced from examining eachof the main cracks on the peened specimen fracture surfaces. The crack growth curves areconsidered to be reasonable representations of the rate of cracking of the largest crack in each ofthese specimens. These curves do not contain the flaw depth, which was also measured duringthis examination. Consequently the data adjacent to the origin is distorted with an accelerationrate higher than the true picture. This will be discussed in Section 4. To improve the clarity of therate of growth while the crack is small the data have also been plotted on log crack depth versus15

DSTO-TR-1477life axes. The data points for these graphs, along with the measured flaw depths are contained inAppendix A.Simulated flight hours0 20000 40000 60000 80000 100000 12000076KSIF 194 peened 360MPaKSIF 107 peened 360MPaKSIF 167 peened 360MPaKSIF 152 peened 360MPaKSIF 149 peened 360MPa0.25Simulated flight hours0 20000 40000 60000 80000 100000 120000100.150.21Crack depth mm.430.150.1Crack depth inch.Crack depth mm.0.10.010.001Crack depth inch.210.050.01KSIF 194 peened 360MPaKSIF 107 peened 360MPaKSIF 167 peened 360MPaKSIF 152 peened 360MPaKSIF 149 peened 360MPa0.0001000 50 100 150 200 250 300 350 400IARPO3a + 5ML Spectra0.0010 50 100 150 200 250 300 350 400IARPO3a + 5ML SpectraABFigure 17 The result of the quantitative fractographic examination of cracks in the peened specimens loadedto a peak stress of 360MPa. Plot ‘A’ presents the results plotted on linear/linear axes. Plot ‘B’presents the same data plotted on log/linear axes.0 4000 8000 12000 16000 20000 24000 28000 3200065Simulated flight hoursKSIF 205 390MPa peenKSIF 151 390MPa peenKSIF 140 390MPa PeenKSIF 118 390MPa peenKSIF 116 390MPa peen0.2Simulated flight hours0104000 8000 12000 16000 20000 24000 28000 320000.11Crack depth mm.4320.150.1Crack depth inch.Crack depth mm.0.10.010.001Crack depth inch.10.050.01KSIF 205 390MPa peenKSIF 151 390MPa peenKSIF 140 390MPa PeenKSIF 118 390MPa peenKSIF 116 390MPa peen0.0001000 20 40 60 80 100IARPO3a + 5ML SpectraA0.0010 20 40 60 80 100IARPO3a + 5ML SpectraFigure 18 The result of the quantitative fractographic examination of cracks in the peened specimens loadedto a peak stress of 390MPa.B16

DSTO-TR-1477Simulated flight hours0 2000 4000 6000 8000 10000 120005Simulated flight hours0 2000 4000 6000 8000 10000 12000104KSIF 191 420MPa peenKSIF 193 420MPa peenKSIF 204 420MPa peenKSIF 147 420MPa peenKSIF 142 420MPa peen0.1610.1Crack depth mm.320.120.08Crack depth inchCrack depth mm.0.10.010.001Crack depth inch1000 5 10 15 20 25 30 35 40IARPO3a + 5ML Spectra0.040.01KSIF 191 420MPa peenKSIF 193 420MPa peenKSIF 204 420MPa peenKSIF 147 420MPa peenKSIF 142 420MPa peen0.0010 5 10 15 20 25 30 35 40IARPO3a + 5ML SpectraABFigure 19 The result of the quantitative fractographic examination of cracks in the peened specimens loadedto a peak stress of 420MPa.0.0001Simulated flight hours0 1000 2000 3000 4000 5000 6000 7000 80003.5Simulated flight hours0 1000 2000 3000 4000 5000 6000 7000 80001032.5KSIF 186 450MPa peenKSIF 146 450MPa peenKSIF 164 450MPa peenKSIF 178 450MPa peenKSIF 183 450MPa peen0.120.110.1Crack depth mm.21.50.080.06Crack depth inch.Crack depth mm.0.10.010.001Crack depth inch.10.040.010.5000 5 10 15 20 25IARPO3a + 5ML SpectraA0.02KSIF 186 450MPa peenKSIF 146 450MPa peenKSIF 164 450MPa peenKSIF 178 450MPa peenKSIF 183 450MPa peen0.0010 5 10 15 20 25IARPO3a + 5ML SpectraFigure 20 The result of the quantitative fractographic examination of cracks in the peened specimens loadedto a peak stress of 450MPa.B0.000117

DSTO-TR-14774. Discussion4.1 IntroductionA number of the observations made during this examination will be discussed in the followingSections, and the results are compared with the results of similar testing on specimens with etchedsurfaces (Barter, 2003).4.2 An examination of testing resultsThe results of the testing are presented in Table 5 and Figure 9. The peening resulted in lifeimprovement over the four peak stress levels tested, when compared to the same specimenswhich had been etched (simulating the etching that occurs to F/A-18 structure prior to theapplication of a corrosion protective aluminium alloy coating by an ion vapour depositionmethod, generally known as an IVD coating). This life improvement is variable and is greatest atthe lower stresses and least at the highest stress while the scatter increases with the lower stresslevels. If the range of three log standard deviations above and below the log mean of the results iscompared to the values for the etched specimens at the same stress levels then the LIF (peenedvalues normalised by dividing by the etched log mean values for each of the stress levels) can becalculated. These results are shown graphically in Figure 21. Exponential curves have been fittedto the data sets, although power curves also fit the data very well. Examining these curvesindicates that the LIFs, measured by the three methods pass through one (no improvement) atabout 475MPa, which is very close to the stated yield stress of this material (461MPa). The factthat the curves indicate a loss of peening life extension at about the yield stress might beintuitively expected, although some workers in this field have suggested that ‘shake down’ atstresses below yield would reduce the effectiveness of the cold working at high sub-yield cyclicstress levels. For this spectrum ‘shake down’ does not appear to be significant. From the data it isclear that the LIF depends on the loading spectrum stress (and probably the spectrum form) andshould not be considered to be a constant for the purposes of calculating the effect of a peeningoperation – the life of the detail that is proposed to be peened must be established given theexpected loading spectrum and the local stressing.1000y = 1.29e+06 * e^(-0.025x) R 2 = 0.983y = 7.16e+07 * e^(-0.0334x) R 2 = 0.962y = 2.31e+04 * e^(-0.0165x) R 2 = 0.971y = 574 * e^(-0.00866x) R 2 = 0.999100y = 2.24e+03 * e^(-0.0163x) R= 0.981y = 1.25e+05 * e^(-0.0247x) R= 0.964y = 40.2 * e^(-0.00789x) R= 0.94210010101Peening log meanPeening + 3 logSDPeening -3 LogSDEtched Log meanPeening log mean/etched meanPeening + 3 logSD/etched meanPeening -3 LogSD/etched mean1350 400 450 500Stress Level MPa0.1350 375 400 425 450 475 500Peak stress level MPa.Figure 21 The log mean curves for the etched and peened tests are shown along with the + and – 3 log SDsfor the peened specimens. Dividing the log mean life and the + and – 3 log SDs of the peenedspecimens by the log mean etched specimen lives gives an indication of the ‘life improvementfactor’ for glass bead peening. Exponential curves have been fitted to the data sets and indicate aloss of peening effect at about the yield stress of this material.18

DSTO-TR-14774.2.1 Position of the cracking and crack originsCrack initiations in the peened specimens were all found to be associated with flaws in thesurface, as was the case with etched specimens for the same specimen type and material. Usuallythe origins of the peened specimens consisted of closely spaced multiple initiations along theedges of laps produced during the peening as noted previously. These are the most prevalent typeof flaw in these and other glass bead peened aluminium alloy surfaces, populating the surface, toa greater or lesser degree, depending on the quality of the peening, as was shown by Clayton &Clark, 1988, Sharp et al, 1994 and Sharp & Clark, 2001. Those laps that are most likely to initiatecracking are usually at the base of peening dents. Deep cuts and embedded fractured glass beadswere also noted as crack starters. These flaws also produced multiple origins around their edges.Although multiple local origins initially occur around a flaw, cracks quickly combined to form asingle crack, which progress more rapidly into the material at its centre than its flanks.Although the predominant position of the largest cracks was at or near corner radii, large crackswere found in all positions; sides, edges, close to the corners and away from the corners, generallywithin 10mm of the specimen centre. The corners, or at least close to the corners of the specimenswould appear to be the dominant regions of cracking due to a number of interacting effects: thecorners have less constraint over the flat surfaces as noted before, the probability of folding at thecorners is higher since the peening is bound to be angled to the surface at some time during thepeening process (for example see Figure 22) and over peening is more likely at the corners sincethey receive peening when the flat surfaces are peened and when the radii are peened as aseparate peening operation.Figure 22 An example of a poorly peened corner. Note the laps (arrows) that indicate surface flow awayfrom the corner in both directions (Not from this specimen series).The dominance of a single crack while many other very small cracks were present indicated thatalthough there may be many crack initiations in the peened specimens only a few cracksdominate. This is most likely the result of the very slow crack growth through the residual stressaffected layer; once a crack has passed though, it grows relatively rapidly to failure, dominatingthe damage to the specimen. This results in the specimen failing before most, if any of the othercracks have exited the region of residual stress influence. This led to some very jagged fracturesurfaces (Figure 10B) due to the fast fracture running along the grain orientation rather thanlinking together significant cracks during failure. Fractures similar to this have also been found in7050 thick section plate specimens with a similar grain orientation with machined surfaces,Aktepe, 2000. In these cases the jagged nature of the final fracture was attributed to a lack ofsecondary cracking due to a low population of surface flaws, which led to the final fracturerunning along the grain orientation rather than across the specimen test section.19

DSTO-TR-14774.2.2 Shape of the crack growth curvesAn examination of the crack growth curves indicated that the cracks are initially retarded by theresidual stresses and then transition to a faster rate of acceleration after the cracks pass a depth ofabout 0.2mm. This acceleration never exceeds the unpeened growth acceleration rate as indicatedby the tests reported in Barter, 2003. This agrees with previous studies of the depth of the residualstress layers below the surface of specimens in this material; Sharp and Clark, 2001, Chang et al,2003. It is notable that the balancing tensile residual stresses are not adversely affecting the crackgrowth. The predicted effect of the residual stress is to lower the crack tip K as the compressivestress is applied to the crack surfaces, and then as the tensile stresses begin to be added, theretarding effect decays.For example, the predicted K may be calculated assuming (for the purpose of the calculation) anapplied static stress of 390MPa and the residual stress distribution shown in Figure 4. Using anembedded thumb nail shaped crack solution, in a large plate where K = 1.12σ(πa) 0.5 , the curvesshown in Figure 23 result. In this Figure the results have been plotted on log-log axes since the Ksolution should be a straight line on this presentation (power law). The first notable point is thedecrease in K due to the residual compression, which is then restored to the non-residual stresssituation when the crack enters the residual tension layer. This lowers the K curve but maintains itparallel to the original. This supports the observation that the region of growth through theresidual compression layer generally has a similar form (exponential, although at a lower slope) tothe growth observed in the non-peened specimens or the region of growth after the residualstresses have been passed. The slight increase in the tensile region, and consequent increase in K,has been observed to have no real growth effect whereas some effect might have been expectedfrom the curves. This is considered to be the result of the clamping action of the residualcompressive stress behind the crack tip being a potent method of retarding the potential crackgrowth rate.100K for a constant 390MPa tensile loadadjusted for residual stressK for a constant 390MPa tensile load1000101Interpolated residual stress frompeening plus 390MPa tensile load0.11000.001 0.01 0.1 1a mm.Figure 23 Here the effect on the K on the residual stress (brown line) is shown compared to no effect (redline). The residual stress lowers the K for about 0.07mm of depth in this case (see footnote 1).Note that the change in K due to the residual stress is mostly parallel to the unaffected K.20

DSTO-TR-14774.3 Estimating the effect of the initiating flawsThe examination of the fatigue crack growth of the KS etched specimens revealed that a relativelysimple method could be used to estimate the crack like size of the flaws from which the cracksgrew. This is reported in Barter, 2003. In that case the flaws that started cracking were etch pitsassociated with grain boundaries and etched out inclusions. The method used to analyse theseverity of these flaws was to fit an exponential growth model of the form:a = a 0 e β (N) (1)Where ‘a’ is the crack depth, ‘a 0 ‘is the apparent crack size at the commencement of loading, ‘N’ isthe life and ‘β’ is the slope of the curve, or a measure of the rate of crack growth. This was handledin several ways; one of which was to fit the curve to all the data points and use the zero lifeintersect as the measure of the Estimated Pre-Crack Size (EPS). A more complete discussion of thisprocess as applied previously may be found in Barter, 2003.The flaws that initiated the cracking in the peened specimens varied, although the dominant flawtype was observed to be laps in the surface due to folding during peening (section 3.6.1).Although the lap depth or the depth of any of the other types of flaw found in these specimenswas usually fairly easy to measure, unlike the etch pit flaws, their length varied along with theirposition within the peening dents. This led to uncertainty about the effectiveness of these flaws ininitiating fatigue cracks, so again the use of the QF results to estimate the EPS seemed to bewarranted.The growth rates of the peened specimens were not consistently exponential over their entirecrack depths ruling out the use of all the data to estimate the EPS. Nevertheless, the growth didappear to be the result of two phases of exponential growth connected by a transition phase atabout the depth at which the retarding effect of the peening disappeared (as was predicted byprevious measurements of the typical residual stress such as reported in Wang, 2003).The EPS was estimated using the slope of the crack growth curves, which in essence assumed thatthe resulting EPS is a combination of the size and shape of the flaw, the flaw position and thenumber of initiations that grew together from the flaw, as well as any abnormality (from theexponential growth rate assumed) in the growth at the very start of the cracking. The confidencein the EPS size was highest for those cracks where the majority of the progressions close to theorigin/s were found.To achieve the most reasonable 4 EPS result for these crack growth curves, several processes havebeen investigated. The easiest approach for the fractographer who is interactively examining aparticular crack and the measurements taken from that crack, is to add the measured depth of theflaw to the raw data which has been measured from the flaw-to-fatigue surface interface.Alternatively, varying amounts of depth can be added until a reasonably straight line is achievedon a log depth versus linear life plot of the measurements for the early part of the crack growth.This ‘anticipation-by-eye’ approach allows a ‘first cut’ at the data, which is used to highlightobvious errors in the measurements so that re-assessment of the measurements can be carried outon the specimen in real-time. This helps prevent ‘uninformed’ data correction while re-measuring4 The most reasonable measure of the EPS is still to be fully defined, although it should have some of thefollowing attributes; should not be so subjective such that different workers would produce grosslydifferent measures of it from the same data, be independent of the spectrum or load level used todetermine it, be transferable to different items made of the same material, not be excessivelyconservative or unconservative, nor should it be grossly different to the real flaw size that started thecracking since this may lead to confusion in interpreting some problems.21

DSTO-TR-1477poorly defined areas on the fracture surface. The interactive examination of the crack surface is astrong advantage of this method, while the dependence of the result on the skill and consistencyof the fractographer may be a disadvantage as far as reproducibility is concerned.After the data have been collected and the flaw depth and initial best-fit EPS have beenanticipated other methods were investigated to estimate EPS. These are listed below along withthe two prior types mentioned as measures of the flaw effectiveness. Each method has beendesignated as an EPS Type:Type I.‘Measured flaw depth’. This is usually the depth of the flaw at the point from whichcrack measurement was started, and has a level of uncertainty associated with theflaw shape and where along the edge of the flaw the measurements were startedfrom, as well as the position of the true surface; either the peaks between the peeningdents or the base of the peening dents or somewhere between. (It should be notedthat this is not necessarily the deepest point of the flaw). In some historicalpresentations of this approach, particularly Barter et al, 1991 and Athiniotis et al,1991 the measured flaw size was included in the data and a exponential curve fit wasapplied (equation 1) to the data to determine the a o (zero life intersect). Thisapproach, although not used here is designated as a Type IA EPS.Type II. The measure that was used during the collection of the QF data, which was the‘anticipated EPS’. This measure is subjective and depends on the skill of thefractographer.Type III. Fitting the data to the exponential model (equation 1) using a curve fitting program,and optimising the initial crack depth added to the raw crack size (offset value),checking the ‘goodness of fit’ (R 2 value), and refining the depth added until themaximum R 2 is reached. This method could be used to produce the curve that bestfitted the total data, and gave a zero life intercept value, which was thought to be thebest measure of the two possible EPSs (offset (Type IIIA) or intercept (Type III)). Thismethod would produce an ‘optimised EPS’ for all the data which clearly was notappropriate here due to the non-single exponential growth rate for the cracksgrowing through the peening.Type IV. The data for each curve were divided into two at the middle of the transition Phase(II) between the slower accelerating early Phase I of cracking and the faster latterPhase III. The ‘anticipated EPS’ was included in the early Phase I data in order togive an approximately correct starting point and an exponential curve was fitted toPhase I data ignoring the effect of the small amount of transition data included. Thezero life intersect of the curve a 0 was then used as the ‘fitted EPS’. This process wasdifferent to the process followed in Barter, 2003 (Type III EPS).While the Type III EPS is probably a more refined method for growth curves that areclose to an exponential over their entire life, the amount of data to use in the case ofthe peened slopes needed to be considered carefully since the optimisation routineused in the Type III EPS calculation, when applied to the early Phase I data,occasionally resulted in a runaway of the of the amount added since the curve simplybecomes a more perfect exponential fit with greater additions when the data is notvery exponential in nature. This was particularly the case when the early part of thecurve is not entirely straight (log depth/linear life) such as some of the curves for the360MPa peak stress.22

DSTO-TR-1477Examples of the Type IV process and the formula generated are shown in Figure 24 and Figure 25for one crack from each peak stress level. These values along with the curve fits for the Phase IIIparts of the curves will be presented and discussed in the following section. Although this methodwhen applied to the peened specimens would probably over estimate the effectiveness of a flawconsidering the variable residual stress (particularly the small region near the surface where theresidual stresses are less intense, see Figure 23), it does give a consistent picture that may betransferable to other spectra for this material.The crack growth curves with corrections for the measured flaw depth (Type I EPS) and theestimated EPS Type II are shown in Figure 26 to Figure 29 A & B. Table 6 gives the Type II EPSand flaw depths along with the means of these values. The Type IV EPS calculations will bepresented in the following section.0 10000 20000 30000 4000010KSIF 149 lowKSIF 149 highSFHy = 0.11611 * e^(0.011033x) R= 0.97153SFH0 5000 10000 15000 20000 2500010KSIF 140 lowKSIF 140 highy = 0.03347 * e^(0.041935x) R= 0.99540.1y = 1.1892e-09 * e^(0.17481x) R= 0.927170.11y = 8.5323e-06 * e^(0.18776x) R= 0.9974110.010.10.010.0010.10 20 40 60 80 100 120 140IARPO3a + 5ML Spectra0.010 10 20 30 40 50 60 70 80IARPO3a + 5ML SpectraFigure 24 Examples of the generation of Phase I and Phase III growth slopes and the Type IV EPS valuesfor the 360 and 290MPa peak stress levels.SFH0 2000 4000 6000 8000 1000010KSIF 191 lowKSIF 191 highSFH0 1000 2000 3000 4000 5000 6000 7000 800010KSIF 146 lowKSIF 146 high1y = 0.043288 * e^(0.089068x) R= 0.98739y = 7.2653e-05 * e^(0.33012x) R= 0.99410.11y = 0.049993 * e^(0.1354x) R= 0.98559y = 0.00028305 * e^(0.4503x) R= 0.944750.10.010.010.10.10.0010.0010.010 5 10 15 20 25 30 35IARPO3a + 5ML Spectra0.010 5 10 15 20 25IARPO3a + 5ML SpectraFigure 25 Examples of the generation of Phase I and Phase III growth slopes and the Type IV EPS valuesfor the 420 and 450MPa peak stress levels.23

DSTO-TR-1477Simulated flight hours0 20000 40000 60000 80000 100000 12000010Simulated flight hours0 20000 40000 60000 80000 100000 120000100.10.111Crack depth mm.0.10.010.001Crack depth inch.Crack depth mm.0.10.010.001Crack depth inch.0.010.01KSIF 194 peened 360MPa + flaw depthKSIF 107 peened 360MPa + flaw depthKSIF 149 peened 360MPa + flaw depthKSIF 152 peened 360MPa + flaw depthKSIF 167 peened 360MPa + flaw depth0.0010 50 100 150 200 250 300 350 400IARPO3a + 5ML SpectraA0.0001KSIF 194 peened 360MPa + EPSKSIF 107 peened 360MPa + EPSKSIF 149 peened 360MPa + EPSKSIF 152 peened 360MPa + EPSKSIF 167 peened 360MPa + EPS0.0010 50 100 150 200 250 300 350 400IARPO3a + 5ML SpectraFigure 26 Crack growth results for the 360MPa peened specimens where the data includes the measuredflaw depth; A and Type II EPS; B.B0.0001Simulated flight hours0104000 8000 12000 16000 20000 24000 28000 32000Simulated flight hours0104000 8000 12000 16000 20000 24000 28000 320000.10.1Crack depth mm.10.01Crack depth inch.Crack depth mm.10.01Crack depth inch.0.10.1KSIF 140 390MPa Peen + flaw depthKSIF 116 390MPa peen + flaw depthKSIF 118 390MPa peen + flaw depthKSIF 205 390MPa peen + flaw depthKSIF 151 390MPa peen + flaw depth0.010 20 40 60 80 100IARPO3a + 5ML SpectraA0.001KSIF 140 390MPa Peen + EPSKSIF 116 390MPa peen + EPSKSIF 118 390MPa peen + EPSKSIF 205 390MPa peen + EPSKSIF 151 390MPa peen + EPS0.010 20 40 60 80 100IARPO3a + 5ML SpectraFigure 27 Crack growth results for the 390MPa peened specimens where the data includes the measuredflaw depth; A and Type II EPS; B.B0.00124

DSTO-TR-1477Simulated flight hours0 2000 4000 6000 8000 10000 1200010Simulated flight hours0 2000 4000 6000 8000 10000 12000100.10.111Crack depth mm.0.10.01Crack depth inch.Crack depth mm.0.10.01Crack depth inch.KSIF 142 420MPa peen + flaw depthKSIF 147 420MPa peen + flaw depthKSIF 204 420MPa peen + flaw depthKSIF 193 420MPa peen + flaw depthKSIF 191 420MPa peen + flaw depth0.010 5 10 15 20 25 30 35 40IARPO3a + 5ML SpectraA0.001KSIF 142 420MPa peen + EPSKSIF 147 420MPa peen + EPSKSIF 204 420MPa peen + EPSKSIF 193 420MPa peen + EPSKSIF 191 420MPa peen + EPS0.010 5 10 15 20 25 30 35 40IARPO3a + 5ML SpectraFigure 28 Crack growth results for the 420MPa peened specimens where the data includes the measuredflaw depth; A and Type II EPS; B.B0.001Simulated flight hours0 1000 2000 3000 4000 5000 6000 7000 800010Simulated flight hours0 1000 2000 3000 4000 5000 6000 7000 8000100.10.111Crack depth mm.0.10.010.001Crack depth inch.Crack depth mm.0.10.010.001Crack depth inch.0.01KSIF 186 450MPa peen + flaw depthKSIF 183 450MPa peen + flaw depthKSIF 178 450MPa peen + flaw depthKSIF 164 450MPa peen + flaw depthKSIF 146 450MPa peen + flaw depth0.0010 5 10 15 20 25IARPO3a + 5ML SpectraA0.00010.01KSIF 186 450MPa peen + EPSKSIF 183 450MPa peen + EPSKSIF 178 450MPa peen + EPSKSIF 164 450MPa peen + EPSKSIF 146 450MPa peen + EPS0.0010 5 10 15 20 25IARPO3a + 5ML SpectraFigure 29 Crack growth results for the 450MPa peened specimens where the data includes the measuredflaw depth; A and Type II EPS; B.B0.000125

DSTO-TR-1477in the appropriate regions of a component. Clearly the number of initiation sites must berelatively large for a consistent measure of the crack growth to be made – the fastest cracks need tobe included in a data set in order to be confident of a prediction, or a distribution of the crackgrowth rates for the material must be known. The optimum specimen size for a particular surfacecondition is as yet unknown, although for etching it would be expected to be considerably smallerthan for peening since etching produces many etch pits that grow similar sized cracks, Barter,2003, while it remains unclear as to the density of the peening flaws in these specimens since mostof the secondary cracks are very much smaller than the lead cracks.The mean slopes of the KS etched specimens, normalised to an EPS of 0.01mm were calculated andthe results are presented graphed, for one set of etched curves (as an example) and combined forall stress levels tested in Figure 30. The EPS parameters for the four stress levels (unnormalised),equivalent to the stress levels tested for the peened specimens are presented in Table 7.Simulated flight hours0 800 1600 2400 3200 4000 4800 5600 640010Simulated flight hours0 5000 10000 15000 20000100.10.111Crack depth mm0.10.01KSIF 105 420MPa + EPSKSIF 121 420MPa + EPSKSIF 136 420MPa + EPSKSIF 171 420MPa + EPSKSIF 173 420MPa + EPSKSIF 180 420MPa + EPS420MPa + EPS mean slop0.0010 5 10 15 20IARPO3a + 5ML Spectra0.010.001Crack depth inch0.00010.10.01270MPa + EPS mean slop300MPa + EPS mean slop330MPa + EPS mean slop360MPa + EPS mean slop390MPa + EPS mean slop420MPa + EPS mean slop450MPa + EPS mean slop0.0010 10 20 30 40 50 60 70IARPO3a + 5ML SpectraFigure 30 Mean crack growth curves for one of the families of etched specimen results from Barter, 2003.The second graph shows the mean curves for each stress level in the etched specimens.Crack depth mm0.010.0010.0001Crack depth inchThe same process was followed for the peened specimens assuming that the crack growth curveshave only two distinct exponential phases of growth as discussed above The parameters for theseslopes are shown in Table 8. Figure 31 and Figure 32 show the comparison of the mean slopes forthe etched specimens and the peened specimens when the EPS is normalised to 0.01mm. It may benoted that the steep part of the peened specimen curves produces a slope that is very close to theslope of the etched specimens, at the same applied stress level, although always lagging the etchedspecimens. This indicates that the effect of the peening is mostly confined to the small region nearthe surface, although even after this region is left behind the cracking proceeds with someretardation evident, possibly due to the surface edges of the cracking still being influenced by thepeening.27

DSTO-TR-1477Table 7Etched specimen calculated growth rate and EPS parametersPeakstresslevel MPaSpecimenNo.Type III EPSa 0(Optimised)Type III slopeβ(Optimised)Average βUsing TypeIII EPS360 KSIF100 0.0271 0.224 0.203KSIF110 0.0253 0.182KSIF129 0.0069 0.242KSIF162 0.026 0.168KSIF198 0.0209 0.197390 KSIF165 0.0085 0.305 0.256KSIF170 0.0228 0.281KSIF181 0.0139 0.266KSIF182 0.0213 0.226KSIF207 0.0394 0.201420 KSIF105 0.0177 0.365 0.380KSIF121 0.0076 0.415KSIF136 0.0129 0.339KSIF171 0.0061 0.392KSIF173 0.0073 0.363KSIF180 0.0059 0.392450 KSID101 0.011 0.489 0.458KSIF103 0.0111 0.408KSIF108 0.0092 0.5KSIF109 0.0207 0.41KSIF126 0.0093 0.482Table 8PeakStressMPa360MPalowerslope)390MPa(lowerslope)420MPa(lowerslope)450MPa(lowerslope)360MPa(higherslope)Calculated growth and EPS parameters for the peened specimensSpecimenNo.Type IV EPSa 0 (mm)Type IVβ (slope)Measureof fit: R 2KSIF 194 0.0651 0.0072 0.944KSIF 107 0.0135 0.0089 0.987KSIF 149 0.1161 0.0110 0.972KSIF 152 0.0095 0.0207 0.999KSIF 167 0.0541 0.0125 0.994KSIF 140 0.0335 0.0419 0.995KSIF 205 0.0711 0.0392 0.994KSIF 118 0.0212 0.0418 0.992KSIF 116 0.0479 0.0377 0.980KSIF 151 0.0261 0.0460 0.986KSIF 142 0.0479 0.1030 0.993KSIF 204 0.0349 0.0838 0.993KSIF 147 0.0270 0.0856 0.985KSIF 193 0.0318 0.0835 0.991KSIF 191 0.0433 0.0891 0.987KSIF 183 0.0309 0.1582 0.998KSIF 178 0.0130 0.3395 0.997KSIF 186 0.0285 0.1651 0.994KSIF 164 0.0110 0.2006 0.996KSIF 146 0.0500 0.1354 0.986KSIF 194 0.0805 0.908KSIF 107 0.1666 0.978KSIF 149 0.1748 0.927KSIF 152 0.0948 0.975KSIF 167 0.1224 0.903log meanof a 0 (mm)averageof βSD oflog a 00.0350 0.0121 0.4670.0363 0.0413 0.2100.0362 0.089 0.1000.0227 0.1998 0.279Average Maxcrack depth(failure) mm.0.1278 5.3828

DSTO-TR-1477390MPa(higherslope)420MPa(higherslope)450MPa(higherslope)KSIF 140 0.1878 0.997KSIF 205 0.2280 0.988KSIF 118 0.2149 0.997KSIF 116 0.2462 0.980KSIF 151 0.2011 0.998KSIF 142 0.3729 0.999KSIF 204 0.2684 0.995KSIF 147 0.2815 0.980KSIF 193 0.2437 0.995KSIF 191 0.3301 0.994KSIF 183 0.3898 0.998KSIF 178 0.013 0.3395 0.989KSIF 186 0.3739 0.996KSIF 164 0.4427 0.998KSIF 146 0.4503 0.9450.2156 4.150.2993 3.630.3993 2.47Note: The a 0 values for the parts of the curve that have the high slope have not been reported since they aremeaningless because the higher slope only occurred after the cracking had grown through the peenedlayer. The exception is the crack measured in Specimen KSIF 178 which did not appear to have beeninfluenced by the peening, suggesting the peening at the origin of this crack was not effective.Simulated flight hours100 40000 80000 120000 1600000.11Crack depth mm0.10.01360MPa peened lower slop390MPa peened lower slop420MPa peened lower slop450MPa peened lower slop360MPa peened high slop390MPa peened high slop420MPa peened high slop450MPa peened high slop0.010.0010.0001Crack depth inch0.0010 100 200 300 400 500 600IARPO3a + 5ML SpectraFigure 31 Mean crack growth slopes of the peened specimens under various peak stresses split into tworegions: the low initial slope associated with the residual stress retardation and the higher finalslope.29

DSTO-TR-1477Simulated flight hours0 40000 80000 120000 16000010Simulated flight hours0 8000 16000 24000 32000 40000 48000100.10.111Crack depth mm0.10.010.001Crack depth inchCrack depth mm0.10.010.001Crack depth inch0.010.01360MPa etched slop360MPa peened lower slop360MPa peened high slop0.0001390MPa etched slop390MPa peened lower slop390MPa peened high slop0.00010.0010 100 200 300 400 500 600IARPO3a + 5ML Spectra0.0010 50 100 150IARPO3a + 5ML SpectraSimulated flight hours0 5000 10000 15000 2000010Simulated flight hours0 2000 4000 6000 80001010.110.1Crack depth mm0.10.010.001Crack depth inchCrack depth mm0.10.010.001Crack depth inch0.010.01420MPa etched slop420MPa peened lower slop420MPa peened high slop0.0001450MPa etched slop450MPa peened lower slop450MPa peened high slop0.00010.0010 10 20 30 40 50 60 700.0010 5 10 15 20 25 30IARPO3a + 5ML SpectraIARPO3a + 5ML SpectraFigure 32The slopes from the peened and etched crack growth data showing the similar slope of the etchedcrack growth with the final slope of the peened specimens.4.5 Comparison of EPS resultsThe EPS results for each EPS Type discussed above are summarised in Table 9. These results areplotted against each other in Figure 33 along with the cumulative probability curves. Thecomparison graphs show that although most of the initial flaws had EPS’s similar to theirmeasured depths there were a considerable number of flaws where this was not the case. ThoseEPS’s higher than the line on the ‘Type II verus Type I’ and ‘Type IV versus Type I’ plots wereusually associated with long shallow flaws and those lower on these graphs, with short deep flaws(which, since the long shallow laps were the dominant flaw in the peened specimens, resulted infew points significantly below the lines). The Type II and Type IV EPS values appear to be verysimilar, although in most cases the Type IV EPS values are larger than the Type II EPS values andas such are considered a more conservative measure of the EPS, notwithstanding theacknowledged level of subjectivity that still remains in their derivation.30

DSTO-TR-1477Table 9Specimen No.A comparison between the different EPS values determined for the glass bead peening ofthe 7050T7451 thick plate studied hereType I EPS(flaw depth)mmType II EPSmm.Type IV EPSmm.% differencebetween Type Iand II EPS% differencebetween Type IIand I VEPS% differencebetween TypeI and IV EPSKSIF 194 0.0884 0.080.06511123 36KSIF 107 0.0064 0.0080.0135-20-41 -53KSIF 149 0.0292 0.10.1161-71-14 -75KSIF 152 0.0092 0.010.0095-85 -3KSIF 167 0.0048 0.050.0541-90-8 -91KSIF 140 0.012 0.030.0335-60-10 -64KSIF 205 0.0324 0.070.0711-54-2 -54KSIF 118 0.0188 0.020.0212-6-6 -11KSIF 116 0.036 0.0450.0479-20-6 -25KSIF 151 0.0368 0.030.02612315 41KSIF 142 0.056 0.050.0479124 17KSIF 204 0.0168 0.030.0349-44-14 -52KSIF 147 0.0308 0.030.027311 14KSIF 193 0.0284 0.030.0318-5-6 -11KSIF 191 0.0208 0.040.0433-48-8 -52KSIF 183 0.028 0.03 0.0309 -7-3 -9KSIF 178 0.013 0.01 0.013 30-23 0KSIF 186 0.0392 0.03 0.0285 315 38KSIF 164 0.0076 0.01 0.011 -24-9 -31KSIF 146 0.0248 0.05 0.05 -500 -50Average %difference -20 -4 -22Also included in Table 9 are the percentage differences between the different measures of the EPS.Some noteworthy outcomes of this comparison are that the lowest stress level results appear to bethe least consistent from measure to measure and that generally the Type II and IV results are verysimilar as would be expected since the Type II results are incorporated in the calculation of theType IV results. It is also noted that the Type II and IV results had about the same difference withthe measured flaw depth (Type I).The cumulative probability curves for these data suggest a log normal distribution will give agood fit as shown by the fitted exponential curves on the final graph in Figure 33. Using thefunction used to generate these curve fits it is possible to predict the size of an out-lying flaw(large flaw of low probability). For instance the 1 in 1000 flaw for the three measures of the EPSare listed in Table 10 along with the lognormal mean and standard deviation.Table 10The 1 in 1000 EPS values predicted for each EPS Type given that the distributions areLognormal. Included are the means and log standard deviations for are of the distributions.Measure of flaw Type Estimated EPS Lognormal Lognormal standardmean mm. deviation SD log (log)I 0.206mm 0.021 0.32II 0.255mm 0.030 0.31IV 0.236mm 0.032 0.2831

DSTO-TR-14770.120.120.10.10.080.080.060.060.040.040.020.0200 0.02 0.04 0.06 0.08 0.1Type I EPS (measured flaw depth) mm.0.1200 0.02 0.04 0.06 0.08 0.1 0.12Type II EPS mm.99.9999.90.1990.080.060.040.0200 0.02 0.04 0.06 0.08 0.1Type I EPS (measured flaw depth) mm.959080705030201051.1Type I EPS (measured flaw depth) mm.Type II EPS mm.Type IV EPS mm..010.001 0.01 0.1 1EPS flaw depth mm.Figure 33Type II EPS plotted against the measured initial flaw depth (Type I EPS). Type VI EPS plottedagainst Type I EPS and the Type IV EPS plotted against Type II EPS. The line in each graphindicates the one-to-one ratio. Finally the cumulative probability plot for the three types of EPSis shown with the best fit curves.An examination of the mean and SD log data given in Table 10 along with the plotted data in thefinal graph of Figure 33 indicates that although the Type II EPS values are on average smaller thanthe Type IV values, the lower Type IV SD log values results in the small value at the 1 in 1000 EPSsize, which is not all that different to the value predicted by the measurement of the flaw depth. Itmay be that with sufficient data, for the purposes of prediction, the use of the flaw depth in thesepeened specimens, and possibly peened F/A-18 structure, may be sufficient for the prediction of aparts life when considering the 1 in 1000 life point.The size of the 1 in 1000 EPSs, although an ‘imaginary’ measure and not an actual crack size,appears to be about the depth where the peening ceases to be fully effective and on these groundsit may be predicted that there will be instances where complex shaped aircraft components whichhave been ‘repaired’ by peening will not receive the expected life extension, and therefore be a32

DSTO-TR-1477considerable concern to aircraft structural integrity. This is borne out by the case where peeningdid not appear to be effective in the KSIF178 specimen, as well as previous experience with theFT55 and the FT488/2 full-scale and component tests (Barter, 1996, Athiniotis et. al, 2003). This,while pessimistic about the practical advantage of glass bead peening for aircraft repair, does notconsider ceramic bead peening which has been shown (Sharp & Clark, 2001) to produce surfaceswith a reduced scatter in their fatigue lives, and thus may produce a considerably smaller 1 in 1000EPS than for glass bead peening.5. ConclusionA series of fatigue coupon tests were carried out to investigate the effect of glass bead peening onF/A-18 aircraft 7050T7451 components. This type of surface treatment has been used as a repair toextend the life of many critical regions on several of the fracture critical bulkheads of the aircraft.These tests produced the following results:1. The peening had in all cases (including KSIF 178) resulted in an extension of the specimenlife when compared to the etched specimens from Barter, 2003.2. The peened specimens gave a scatter on life dependent on the peak stress level. The lowerthe stress the higher the scatter.3. The Life Improvement Factor for the peened specimens compared to the etched specimensreported in Barter, 2003, varied exponentially with the peak stress level reducing to noapparent life improvement when the spectrum peak stress level was about the yield stressof the material.4. The crack growth curves for all specimens had growth rates that were well represented bytwo phases of exponential growth, with a early stage of retarded growth as the crack grewthrough the residual stress layer and a faster growth acceleration phase almost as fast asthe growth of cracking without the effect of peening. A small beneficial effect on thegrowth of the faster phase was noted, although it would need to be further investigated toconfirm this as a consistent effect.5. All lead cracks had initiated in the first applied block of loading. Many other secondarycracks existed in the specimens at failure although most of these were very small.6. Several different measures of the effectiveness of the crack initiating defects in the peeningwere evaluated. The Log normal distributions of the measures appeared to be similarwhen used to predict a large flaw of low probability.In general the quantitative fractography gave valuable information about the nature of the crackgrowth. This allowed a better understanding of the life results and the way that peening improvedthe fatigue life of these specimens.The estimated pre-crack size (EPS) values calculated in this series of tests would appear to showpromise as the starting point for a fracture mechanics based prediction of life in peened 7050T7451alloy parts.6. AcknowledgementsThe author gratefully acknowledges the work of Chris Niessen and Brian Jones of DSTO-PSL forspecimen preparation and performing the fatigue tests.33

DSTO-TR-14777. BibliographyAktepe, B. and Dickinson, T., Coupon Tests to Determine the Influence of Buffet Detected at F/A-18 Wing Root. DSTO-TR-0924, Department of Defence, Defence Science and TechnologyOrganisation, 2000.Athiniotis, N., Barter, S. A. and Clark, G., Summary of Fatigue Cracking in RAAF Macchi MB326HWing Spars. ARL-MAT-TR-0747, AR-010-700, Department of Defence, Defence Science andTechnology Organisation. 1999.Athiniotis, N., Barter, S. A., Bohret, D.D., Green, A.J., Houston M.I. and Stimson, M.G., FinalReport for the Component Fatigue Test of a F/A-18 Centre Fuselage FS488 Bulkhead - FT488/2,DSTO-TR-0948, Department of Defence, Defence Science and Technology Organisation. 2003.Barter, S. A., FS488 Bulkhead Fracture Surface Preliminary Fractographic Examination, DefectAssessment and Failure Analysis Report M45/90. Department of Defence, Defence Science andTechnology Organisation. 1990a.Barter, S. A., Athiniotis, N. and Lambianidis L., Examination of the Microstructure of SeveralSamples of 7050 Aluminium Alloy. ARL-MAT-TM-403, AR-006-113, Department of Defence,Defence Science and Technology Organisation. 1990b.Barter, S. A., The use of the Conventional Optical Microscope for Quantitative FractographyMetallography and Mineralogy Conference, Institute of Metals and Materials Australasia Limited.1991.Barter, S. A., Bishop, B. and Clark, G., Defect Assessment on F/A-18 488 Bulkhead Tested at ARL.Aircraft Material Report 125, AR-006-618, Department of Defence, Defence Science andTechnology Organisation. 1991.Barter, S. A., Inspection of Cracking in the Starboard Forward Flange 6 inch Radius of FT488/2.DSTO-DDP-0191, Airframes and Engines Division, Department of Defence, Defence Science andTechnology Organisation. 1996.Barter, S. A., Fractographic Inspection of the Cracking in FT488/2 Removed at Program 79. DSTO-TN-0170. Airframes and Engines Division, Department of Defence, Defence Science andTechnology Organisation. 1998.Barter, S., Molent, L., Sharp, K. and Clark, G., Repair and Life Assessment of Critical FatigueDamaged Aluminium Alloy Structure Using a Peening Rework Method. In the proceedings of theUSAF ASIP Conference, San Antonio USA. 2000.Barter, S. A. and Price, J., Effect of Surface Preparation Treatments on Fatigue Life of 7050-Aluminium Alloy. Proceedings of the Structural Integrity and Fracture 2000 Symposium, pp140-153. 2000.34

DSTO-TR-1477Barter, S. A., Fatigue Crack Growth in 7050T7451 Aluminium Alloy Thick Section Plate with aSurface Condition Simulating Some Regions of F/A-18 Structure. DSTO-TR-1458, Department ofDefence, Defence Science and Technology Organisation. 2003.Clayton, J. Q. and Clark, G., The effect of Steel Shot and Glass Bead Peening Treatments on theFatigue Resistance of 7050-T76351 Aluminium Alloy, Proc. Aust. Ftact. Group, Fracture Mechanicsin Engineering Practice, Melbourne. pp44-51. 1988.Clark, G., Modelling Residual Stresses and Fatigue Crack Growth at Cold-Expanded FastenerHoles. Fatigue Fract. Engng Mater. Struct. Vol. 14, No. 5, pp. 379-589, 1991.Clark, G. and Clayton, J. Q., Effectiveness of Peening Treatments in Improving Fatigue Resistanceof 7050 Aluminium Alloy under Constant Amplitude and Spectrum Loading, Aust. Surface EngngConf., University of Sth. Australia, 1991.Miller, K., The Short Crack problem., Fatigue Fract. Engng Mater. Struct. Vol. 5, No. 3, pp. 223-232.1982.Molent, L., Ogden, R. and Pell, R., F/A-18 FS488 Bulkhead Fatigue Coupon Test Program. DSTO-TR-0941, Airframes and Engines Division, Department of Defence, Defence Science andTechnology Organisation. 2000.Sharp, P. K. and Clark, G., The Effects of Peening on the Fatigue Life of 7050 Aluminium Alloy.DSTO-RR-0208, AR N0. 011-795, Department of Defence, Defence Science and TechnologyOrganisation. 2001.Sharp, P. K, Byrnes, R. and Clark, G., Examination of 7050 Fatigue Crack Growth Data and itsEffect on Life Prediction. DSTO-TR-0729, AR-010-648, Department of Defence, Defence Scienceand Technology Organisation. 1998.Sharp, P. K, Clayton, J. Q. and Clark, G., The Fatigue Resistance of Peened 7050-T7451 AluminiumAlloy – Repair and Re-Treatment of Component Surface, Fatigue Fract. Eng Mater Structure. Vol17, No3, pp243-252. 1994.Simpson, D.L., Molent, L., Landry, N., Roussel, J., Graham, A.D. and Schmidt, N., “The Canadianand Australian F/A-18 International Follow-On Structural Test Project”, Proc. ICAS 2002Congress, Toronto, Canada. 2002.Wang, C. H., Barter, S. A. and Liu, Q., A Closure Model to Crack Growth Under Large-ScaleYielding and Through Residual Stress Fields. Journal of Engineering Material and Technology,Vol. 125. 2003.35

DSTO-TR-1477Appendix A: QF measurements for the Peened SpecimensTable A-1. Results of the quantitative fractography on KSIF 194, peak stress 360MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0884 0.0865.5 21282 0.0256 0.114 0.105666.5 21607 0.026 0.1144 0.10668.5 22257 0.0268 0.1152 0.106870.5 22907 0.0276 0.116 0.107672.5 23557 0.0284 0.1168 0.108473.5 23882 0.0292 0.1176 0.109275.5 24531 0.03 0.1184 0.1177.5 25181 0.0308 0.1192 0.110879.5 25831 0.0316 0.12 0.111680.5 26156 0.0328 0.1212 0.112881.5 26481 0.0336 0.122 0.113682.5 26806 0.0344 0.1228 0.114483.5 27131 0.0352 0.1236 0.115284.5 27456 0.0356 0.124 0.115685.5 27781 0.0364 0.1248 0.116486.5 28106 0.0372 0.1256 0.117287.5 28431 0.038 0.1264 0.11888.5 28755 0.0384 0.1268 0.118489.5 29080 0.0396 0.128 0.119690.5 29405 0.0404 0.1288 0.120491.5 29730 0.0416 0.13 0.121692.5 30055 0.0424 0.1308 0.122493.5 30380 0.0432 0.1316 0.123294.5 30705 0.0444 0.1328 0.124495.5 31030 0.0456 0.134 0.125696.5 31355 0.0468 0.1352 0.126897.5 31680 0.0476 0.136 0.127698.5 32005 0.0484 0.1368 0.128499.5 32330 0.05 0.1384 0.13100.5 32654 0.0512 0.1396 0.1312101.5 32979 0.0524 0.1408 0.1324102.5 33304 0.0536 0.142 0.1336103.5 33629 0.0552 0.1436 0.1352104.5 33954 0.0564 0.1448 0.1364105.5 34279 0.0573 0.1457 0.1373106.5 34604 0.0589 0.1473 0.1389107.5 34929 0.0606 0.149 0.1406108.5 35254 0.0622 0.1506 0.1422109.5 35579 0.0648 0.1532 0.1448110.5 35904 0.0672 0.1556 0.1472111.5 36229 0.0697 0.1581 0.1497112.5 36553 0.0721 0.1605 0.1521113.5 36878 0.0743 0.1627 0.1543114.5 37203 0.0772 0.1656 0.1572115.5 37528 0.0805 0.1689 0.1605116.5 37853 0.0831 0.1715 0.1631117.5 38178 0.0886 0.177 0.1686118.5 38503 0.0937 0.1821 0.1737119.5 38828 0.1003 0.1887 0.1803120.5 39153 0.1072 0.1956 0.1872121.5 39478 0.11 0.1984 0.19122.5 39803 0.1193 0.2077 0.1993123.5 40128 0.1279 0.2163 0.2079124.5 40453 0.1374 0.2258 0.2174125.5 40777 0.1476 0.236 0.227636

DSTO-TR-1477126.5 41102 0.1638 0.2522 0.2438127.5 41427 0.1781 0.2665 0.2581128.5 41752 0.1942 0.2826 0.2742129.5 42077 0.2124 0.3008 0.2924130.5 42402 0.2275 0.3159 0.3075131.5 42727 0.253 0.3414 0.333132.5 43052 0.2636 0.352 0.3436133.5 43377 0.2826 0.371 0.3626134.5 43702 0.3103 0.3987 0.3903135.5 44027 0.3313 0.4197 0.4113136.5 44352 0.3623 0.4507 0.4423137.5 44676 0.3973 0.4857 0.4773138.5 45001 0.4284 0.5168 0.5084139.5 45326 0.4709 0.5593 0.5509140.5 45651 0.5195 0.6079 0.5995141.5 45976 0.5801 0.6685 0.6601143.5 46626 0.6817 0.7701 0.7617144.5 46951 0.7435 0.8319 0.8235145.5 47276 0.8181 0.9065 0.8981146.5 47601 0.8897 0.9781 0.9697147.5 47926 0.9878 1.0762 1.0678148.5 48251 1.1356 1.224 1.2156149.5 48576 1.2591 1.3475 1.3391150.5 48900 1.3887 1.4771 1.4687152.5 49550 1.7064 1.7948 1.7864153.5 49875 2.0829 2.1713 2.1629154.5 50200 2.6026 2.691 2.6826155.5 50525 3.4098 3.4982 3.4898156.5 50850 5.4922 5.5806 5.5722Table A-2. Results of the quantitative fractography on KSIF 107, peak stress 360MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0064 0.008144.5 46951 0.0314 0.0378 0.0394145.5 47276 0.0317 0.0381 0.0397156.5 50850 0.0412 0.0476 0.0492157.5 51175 0.042 0.0484 0.05158.5 51500 0.0431 0.0495 0.0511159.5 51825 0.0438 0.0502 0.0518160.5 52150 0.0449 0.0513 0.0529161.5 52475 0.0453 0.0517 0.0533162.5 52800 0.0464 0.0528 0.0544163.5 53124 0.0468 0.0532 0.0548164.5 53449 0.0476 0.054 0.0556165.5 53774 0.0491 0.0555 0.0571166.5 54099 0.0498 0.0562 0.0578167.5 54424 0.0506 0.057 0.0586168.5 54749 0.051 0.0574 0.059169.5 55074 0.0521 0.0585 0.0601170.5 55399 0.0525 0.0589 0.0605171.5 55724 0.0533 0.0597 0.0613172.5 56049 0.054 0.0604 0.062173.5 56374 0.0552 0.0616 0.0632174.5 56699 0.0559 0.0623 0.0639175.5 57023 0.0567 0.0631 0.0647176.5 57348 0.0579 0.0643 0.0659177.5 57673 0.0586 0.065 0.0666178.5 57998 0.0594 0.0658 0.0674179.5 58323 0.0606 0.067 0.0686180.5 58648 0.0617 0.0681 0.0697181.5 58973 0.0625 0.0689 0.070537

DSTO-TR-1477182.5 59298 0.0633 0.0697 0.0713183.5 59623 0.064 0.0704 0.072184.5 59948 0.0648 0.0712 0.0728185.5 60273 0.0656 0.072 0.0736186.5 60598 0.0664 0.0728 0.0744187.5 60922 0.0671 0.0735 0.0751188.5 61247 0.0679 0.0743 0.0759189.5 61572 0.0683 0.0747 0.0763190.5 61897 0.0691 0.0755 0.0771191.5 62222 0.0699 0.0763 0.0779192.5 62547 0.0706 0.077 0.0786193.5 62872 0.0718 0.0782 0.0798194.5 63197 0.0722 0.0786 0.0802195.5 63522 0.0734 0.0798 0.0814196.5 63847 0.0741 0.0805 0.0821197.5 64172 0.0749 0.0813 0.0829198.5 64497 0.0761 0.0825 0.0841199.5 64822 0.0769 0.0833 0.0849200.5 65146 0.0781 0.0845 0.0861201.5 65471 0.0788 0.0852 0.0868202.5 65796 0.08 0.0864 0.088203.5 66121 0.0808 0.0872 0.0888204.5 66446 0.082 0.0884 0.09205.5 66771 0.0828 0.0892 0.0908206.5 67096 0.0835 0.0899 0.0915207.5 67421 0.0847 0.0911 0.0927208.5 67746 0.0851 0.0915 0.0931209.5 68071 0.0863 0.0927 0.0943210.5 68396 0.0871 0.0935 0.0951211.5 68721 0.0875 0.0939 0.0955212.5 69046 0.0883 0.0947 0.0963213.5 69370 0.089 0.0954 0.097214.5 69695 0.0898 0.0962 0.0978215.5 70020 0.0906 0.097 0.0986216.5 70345 0.0914 0.0978 0.0994217.5 70670 0.0918 0.0982 0.0998218.5 70995 0.093 0.0994 0.101219.5 71320 0.0934 0.0998 0.1014220.5 71645 0.0946 0.101 0.1026221.5 71970 0.0954 0.1018 0.1034222.5 72295 0.0961 0.1025 0.1041223.5 72620 0.0969 0.1033 0.1049224.5 72945 0.0981 0.1045 0.1061225.5 73269 0.0989 0.1053 0.1069226.5 73594 0.0997 0.1061 0.1077227.5 73919 0.1009 0.1073 0.1089228.5 74244 0.1017 0.1081 0.1097229.5 74569 0.1025 0.1089 0.1105230.5 74894 0.1032 0.1096 0.1112231.5 75219 0.104 0.1104 0.112232.5 75544 0.1048 0.1112 0.1128233.5 75869 0.1056 0.112 0.1136234.5 76194 0.1064 0.1128 0.1144235.5 76519 0.1072 0.1136 0.1152236.5 76844 0.108 0.1144 0.116237.5 77169 0.1088 0.1152 0.1168238.5 77493 0.1096 0.116 0.1176239.5 77818 0.1104 0.1168 0.1184240.5 78143 0.1112 0.1176 0.1192241.5 78468 0.1119 0.1183 0.1199242.5 78793 0.1127 0.1191 0.1207243.5 79118 0.1139 0.1203 0.1219244.5 79443 0.1147 0.1211 0.1227245.5 79768 0.1155 0.1219 0.1235246.5 80093 0.1167 0.1231 0.1247247.5 80418 0.1183 0.1247 0.126338

DSTO-TR-1477248.5 80743 0.1195 0.1259 0.1275249.5 81068 0.1203 0.1267 0.1283250.5 81392 0.1211 0.1275 0.1291251.5 81717 0.1223 0.1287 0.1303252.5 82042 0.123 0.1294 0.131253.5 82367 0.1238 0.1302 0.1318254.5 82692 0.1246 0.131 0.1326255.5 83017 0.1254 0.1318 0.1334256.5 83342 0.1262 0.1326 0.1342257.5 83667 0.127 0.1334 0.135258.5 83992 0.1278 0.1342 0.1358259.5 84317 0.1282 0.1346 0.1362260.5 84642 0.1294 0.1358 0.1374261.5 84967 0.1298 0.1362 0.1378262.5 85291 0.131 0.1374 0.139263.5 85616 0.1314 0.1378 0.1394264.5 85941 0.1322 0.1386 0.1402265.5 86266 0.133 0.1394 0.141266.5 86591 0.1342 0.1406 0.1422267.5 86916 0.135 0.1414 0.143268.5 87241 0.1357 0.1421 0.1437269.5 87566 0.1365 0.1429 0.1445270.5 87891 0.1373 0.1437 0.1453271.5 88216 0.1381 0.1445 0.1461272.5 88541 0.1389 0.1453 0.1469273.5 88866 0.1397 0.1461 0.1477274.5 89191 0.1406 0.147 0.1486275.5 89515 0.1414 0.1478 0.1494276.5 89840 0.1422 0.1486 0.1502277.5 90165 0.143 0.1494 0.151278.5 90490 0.1438 0.1502 0.1518279.5 90815 0.1445 0.1509 0.1525280.5 91140 0.1453 0.1517 0.1533281.5 91465 0.1461 0.1525 0.1541282.5 91790 0.1469 0.1533 0.1549283.5 92115 0.1473 0.1537 0.1553284.5 92440 0.1481 0.1545 0.1561285.5 92765 0.1497 0.1561 0.1577286.5 93090 0.1517 0.1581 0.1597287.5 93414 0.1525 0.1589 0.1605288.5 93739 0.1541 0.1605 0.1621289.5 94064 0.1553 0.1617 0.1633290.5 94389 0.1565 0.1629 0.1645291.5 94714 0.1577 0.1641 0.1657292.5 95039 0.1593 0.1657 0.1673293.5 95364 0.1609 0.1673 0.1689294.5 95689 0.1629 0.1693 0.1709295.5 96014 0.1645 0.1709 0.1725296.5 96339 0.1661 0.1725 0.1741297.5 96664 0.1669 0.1733 0.1749298.5 96989 0.1685 0.1749 0.1765299.5 97314 0.1705 0.1769 0.1785300.5 97638 0.1725 0.1789 0.1805301.5 97963 0.1741 0.1805 0.1821302.5 98288 0.1765 0.1829 0.1845303.5 98613 0.1777 0.1841 0.1857304.5 98938 0.1797 0.1861 0.1877305.5 99263 0.1813 0.1877 0.1893306.5 99588 0.1829 0.1893 0.1909307.5 99913 0.1837 0.1901 0.1917308.5 1.00E+05 0.1861 0.1925 0.1941309.5 1.01E+05 0.1885 0.1949 0.1965310.5 1.01E+05 0.1905 0.1969 0.1985311.5 1.01E+05 0.1921 0.1985 0.2001312.5 1.02E+05 0.1944 0.2008 0.2024313.5 1.02E+05 0.1968 0.2032 0.204839

DSTO-TR-1477314.5 1.02E+05 0.1992 0.2056 0.2072315.5 1.03E+05 0.2012 0.2076 0.2092316.5 1.03E+05 0.2036 0.21 0.2116317.5 1.03E+05 0.206 0.2124 0.214318.5 1.03E+05 0.21 0.2164 0.218319.5 1.04E+05 0.2124 0.2188 0.2204320.5 1.04E+05 0.2156 0.222 0.2236321.5 1.04E+05 0.2184 0.2248 0.2264322.5 1.05E+05 0.2212 0.2276 0.2292323.5 1.05E+05 0.2244 0.2308 0.2324324.5 1.05E+05 0.2272 0.2336 0.2352325.5 1.06E+05 0.2304 0.2368 0.2384326.5 1.06E+05 0.234 0.2404 0.242327.5 1.06E+05 0.2376 0.244 0.2456328.5 1.07E+05 0.2416 0.248 0.2496329.5 1.07E+05 0.246 0.2524 0.254330.5 1.07E+05 0.2504 0.2568 0.2584331.5 1.08E+05 0.2568 0.2632 0.2648332.5 1.08E+05 0.262 0.2684 0.27333.5 1.08E+05 0.27 0.2764 0.278334.5 1.09E+05 0.276 0.2824 0.284335.5 1.09E+05 0.2828 0.2892 0.2908336.5 1.09E+05 0.29 0.2964 0.298337.5 1.10E+05 0.3008 0.3072 0.3088338.5 1.10E+05 0.3103 0.3167 0.3183339.5 1.10E+05 0.3191 0.3255 0.3271340.5 1.11E+05 0.3391 0.3455 0.3471341.5 1.11E+05 0.3739 0.3803 0.3819342.5 1.11E+05 0.4051 0.4115 0.4131343.5 1.12E+05 0.4487 0.4551 0.4567344.5 1.12E+05 0.5054 0.5118 0.5134345.5 1.12E+05 0.5754 0.5818 0.5834346.5 1.13E+05 0.6646 0.671 0.6726347.5 1.13E+05 0.7695 0.7759 0.7775348.5 1.13E+05 0.915 0.9214 0.923349.5 1.14E+05 1.1204 1.1268 1.1284350.5 1.14E+05 1.4316 1.438 1.4396351.5 1.14E+05 1.8921 1.8985 1.9001354.5 1.15E+05 3.905 3.9114 3.913Table A-3. Results of the quantitative fractography on KSIF 149, peak stress 360MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0292 0.10.5 162.46 0.002 0.0312 0.1021.5 487.38 0.0044 0.0336 0.10443.5 1137.2 0.0076 0.0368 0.10765.5 1787.1 0.0108 0.04 0.11087.5 2436.9 0.0148 0.044 0.114815.5 5036.3 0.0361 0.0653 0.136116.5 5361.2 0.0393 0.0685 0.139318.5 6011 0.0438 0.073 0.143819.5 6335.9 0.0458 0.075 0.145820.5 6660.9 0.0488 0.078 0.148821.5 6985.8 0.0512 0.0804 0.151222.5 7310.7 0.0536 0.0828 0.153624.5 7960.5 0.0596 0.0888 0.159625.5 8285.5 0.0619 0.0911 0.161926.5 8610.4 0.0639 0.0931 0.163927.5 8935.3 0.0654 0.0946 0.165428.5 9260.2 0.0674 0.0966 0.167429.5 9585.1 0.069 0.0982 0.16940

DSTO-TR-147730.5 9910.1 0.0713 0.1005 0.171331.5 10235 0.0737 0.1029 0.173732.5 10560 0.0756 0.1048 0.175633.5 10885 0.078 0.1072 0.17834.5 11210 0.0804 0.1096 0.180435.5 11535 0.0819 0.1111 0.181936.5 11860 0.0831 0.1123 0.183137.5 12185 0.0847 0.1139 0.184738.5 12509 0.0871 0.1163 0.187140.5 13159 0.093 0.1222 0.19341.5 13484 0.0948 0.124 0.194842.5 13809 0.0964 0.1256 0.196443.5 14134 0.0994 0.1286 0.199444.5 14459 0.1019 0.1311 0.201945.5 14784 0.1035 0.1327 0.203546.5 15109 0.105 0.1342 0.20547.5 15434 0.1068 0.136 0.206848.5 15759 0.1078 0.137 0.207849.5 16084 0.1093 0.1385 0.209350.5 16408 0.111 0.1402 0.21151.5 16733 0.113 0.1422 0.21352.5 17058 0.1148 0.144 0.214853.5 17383 0.1162 0.1454 0.216254.5 17708 0.1173 0.1465 0.217355.5 18033 0.1194 0.1486 0.219456.5 18358 0.1211 0.1503 0.221157.5 18683 0.1232 0.1524 0.223258.5 19008 0.1244 0.1536 0.224459.5 19333 0.1263 0.1555 0.226360.5 19658 0.1279 0.1571 0.227961.5 19983 0.1302 0.1594 0.230262.5 20307 0.1323 0.1615 0.232363.5 20632 0.1335 0.1627 0.233564.5 20957 0.1353 0.1645 0.235365.5 21282 0.1374 0.1666 0.237466.5 21607 0.139 0.1682 0.23967.5 21932 0.1406 0.1698 0.240668.5 22257 0.1433 0.1725 0.243369.5 22582 0.1448 0.174 0.244870.5 22907 0.1468 0.176 0.246871.5 23232 0.1487 0.1779 0.248772.5 23557 0.1497 0.1789 0.249773.5 23882 0.1506 0.1798 0.250674.5 24207 0.1528 0.182 0.252875.5 24531 0.1546 0.1838 0.254676.5 24856 0.1562 0.1854 0.256277.5 25181 0.1578 0.187 0.257878.5 25506 0.1591 0.1883 0.259179.5 25831 0.1615 0.1907 0.261580.5 26156 0.1643 0.1935 0.264381.5 26481 0.1674 0.1966 0.267482.5 26806 0.1691 0.1983 0.269183.5 27131 0.1719 0.2011 0.271984.5 27456 0.1751 0.2043 0.275185.5 27781 0.1782 0.2074 0.278286.5 28106 0.1807 0.2099 0.280787.5 28431 0.1843 0.2135 0.284388.5 28755 0.1872 0.2164 0.287289.5 29080 0.1902 0.2194 0.290290.5 29405 0.1959 0.2251 0.295991.5 29730 0.2004 0.2296 0.300492.5 30055 0.205 0.2342 0.30593.5 30380 0.2089 0.2381 0.308994.5 30705 0.2132 0.2424 0.313295.5 31030 0.2144 0.2436 0.314496.5 31355 0.2198 0.249 0.319841

DSTO-TR-147797.5 31680 0.226 0.2552 0.32698.5 32005 0.2307 0.2599 0.330799.5 32330 0.2369 0.2661 0.3369100.5 32654 0.2436 0.2728 0.3436101.5 32979 0.2502 0.2794 0.3502102.5 33304 0.2572 0.2864 0.3572103.5 33629 0.2631 0.2923 0.3631104.5 33954 0.2713 0.3005 0.3713105.5 34279 0.278 0.3072 0.378106.5 34604 0.291 0.3202 0.391107.5 34929 0.3044 0.3336 0.4044108.5 35254 0.3189 0.3481 0.4189109.5 35579 0.3434 0.3726 0.4434110.5 35904 0.3705 0.3997 0.4705111.5 36229 0.4179 0.4471 0.5179112.5 36553 0.4563 0.4855 0.5563113.5 36878 0.4989 0.5281 0.5989114.5 37203 0.5555 0.5847 0.6555115.5 37528 0.5978 0.627 0.6978116.5 37853 0.6592 0.6884 0.7592117.5 38178 0.7313 0.7605 0.8313118.5 38503 0.844 0.8732 0.944119.5 38828 0.9901 1.0193 1.0901120.5 39153 1.1994 1.2286 1.2994121.5 39478 1.5298 1.559 1.6298122.5 39803 1.9676 1.9968 2.0676123.5 40128 2.6741 2.7033 2.7741124.5 40453 4.0683 4.0975 4.1683125.2 40680 6.3938 6.423 6.4938Table A-4. Results of the quantitative fractography on KSIF 152, peak stress 360MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0092 0.0176.5 24856 0.0371 0.0463 0.047177.5 25181 0.0379 0.0471 0.047978.5 25506 0.0383 0.0475 0.048379.5 25831 0.039 0.0482 0.04980.5 26156 0.0402 0.0494 0.050281.5 26481 0.0406 0.0498 0.050682.5 26806 0.0414 0.0506 0.051483.5 27131 0.0426 0.0518 0.052684.5 27456 0.0434 0.0526 0.053485.5 27781 0.0446 0.0538 0.054686.5 28106 0.0458 0.055 0.055887.5 28431 0.0466 0.0558 0.056688.5 28755 0.0474 0.0566 0.057489.5 29080 0.0486 0.0578 0.058690.5 29405 0.0502 0.0594 0.060291.5 29730 0.051 0.0602 0.06192.5 30055 0.0526 0.0618 0.062693.5 30380 0.0534 0.0626 0.063494.5 30705 0.055 0.0642 0.06595.5 31030 0.0566 0.0658 0.066696.5 31355 0.0582 0.0674 0.068297.5 31680 0.0598 0.069 0.069898.5 32005 0.0614 0.0706 0.071499.5 32330 0.0626 0.0718 0.0726100.5 32654 0.0642 0.0734 0.0742101.5 32979 0.0673 0.0765 0.0773102.5 33304 0.0689 0.0781 0.0789103.5 33629 0.0709 0.0801 0.080942

DSTO-TR-1477104.5 33954 0.0733 0.0825 0.0833105.5 34279 0.0753 0.0845 0.0853106.5 34604 0.0781 0.0873 0.0881107.5 34929 0.0797 0.0889 0.0897108.5 35254 0.0821 0.0913 0.0921109.5 35579 0.0829 0.0921 0.0929110.5 35904 0.0853 0.0945 0.0953111.5 36229 0.0877 0.0969 0.0977112.5 36553 0.0905 0.0997 0.1005113.5 36878 0.0929 0.1021 0.1029114.5 37203 0.0949 0.1041 0.1049115.5 37528 0.0965 0.1057 0.1065116.5 37853 0.0997 0.1089 0.1097117.5 38178 0.1017 0.1109 0.1117118.5 38503 0.1033 0.1125 0.1133119.5 38828 0.1069 0.1161 0.1169120.5 39153 0.1093 0.1185 0.1193121.5 39478 0.1117 0.1209 0.1217122.5 39803 0.1149 0.1241 0.1249123.5 40128 0.1177 0.1269 0.1277124.5 40453 0.1209 0.1301 0.1309125.5 40777 0.1237 0.1329 0.1337126.5 41102 0.1269 0.1361 0.1369127.5 41427 0.1293 0.1385 0.1393128.5 41752 0.1317 0.1409 0.1417129.5 42077 0.1337 0.1429 0.1437130.5 42402 0.1361 0.1453 0.1461131.5 42727 0.1385 0.1477 0.1485132.5 43052 0.1413 0.1505 0.1513133.5 43377 0.1441 0.1533 0.1541134.5 43702 0.1469 0.1561 0.1569135.5 44027 0.1493 0.1585 0.1593136.5 44352 0.1509 0.1601 0.1609137.5 44676 0.1549 0.1641 0.1649138.5 45001 0.1577 0.1669 0.1677139.5 45326 0.1601 0.1693 0.1701140.5 45651 0.1637 0.1729 0.1737141.5 45976 0.1661 0.1753 0.1761142.5 46301 0.1689 0.1781 0.1789143.5 46626 0.1725 0.1817 0.1825144.5 46951 0.1765 0.1857 0.1865145.5 47276 0.1797 0.1889 0.1897146.5 47601 0.1821 0.1913 0.1921147.5 47926 0.1857 0.1949 0.1957148.5 48251 0.1901 0.1993 0.2001149.5 48576 0.1956 0.2048 0.2056150.5 48900 0.1976 0.2068 0.2076151.5 49225 0.2016 0.2108 0.2116152.5 49550 0.2064 0.2156 0.2164153.5 49875 0.2124 0.2216 0.2224154.5 50200 0.2168 0.226 0.2268155.5 50525 0.2244 0.2336 0.2344156.5 50850 0.2304 0.2396 0.2404157.5 51175 0.236 0.2452 0.246158.5 51500 0.2448 0.254 0.2548159.5 51825 0.2536 0.2628 0.2636160.5 52150 0.2636 0.2728 0.2736161.5 52475 0.2752 0.2844 0.2852162.5 52800 0.286 0.2952 0.296163.5 53124 0.2956 0.3048 0.3056164.5 53449 0.3112 0.3204 0.3212165.5 53774 0.3264 0.3356 0.3364166.5 54099 0.3424 0.3516 0.3524167.5 54424 0.3644 0.3736 0.3744168.5 54749 0.384 0.3932 0.394169.5 55074 0.4084 0.4176 0.418443

DSTO-TR-1477170.5 55399 0.4357 0.4449 0.4457171.5 55724 0.4638 0.473 0.4738172.5 56049 0.4943 0.5035 0.5043173.5 56374 0.5254 0.5346 0.5354174.5 56699 0.5675 0.5767 0.5775175.5 57023 0.6112 0.6204 0.6212176.5 57348 0.6573 0.6665 0.6673177.5 57673 0.7208 0.73 0.7308178.5 57998 0.7857 0.7949 0.7957179.5 58323 0.8591 0.8683 0.8691180.5 58648 0.9432 0.9524 0.9532181.5 58973 1.0599 1.0691 1.0699182.5 59298 1.195 1.2042 1.205183.5 59623 1.4078 1.417 1.4178186.5 60598 2.1082 2.1174 2.1182189.5 61572 3.7248 3.734 3.7348190.5 61897 4.5386 4.5478 4.5486191.5 62222 5.0389 5.0481 5.0489192.2 62450 5.504 5.5132 5.514Table A-5. Results of the quantitative fractography on KSIF 167, peak stress 360MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0048 0.055.5 1787.1 0.0092 0.014 0.05926.5 2112 0.0104 0.0152 0.06047.5 2436.9 0.0112 0.016 0.06128.5 2761.8 0.012 0.0168 0.0629.5 3086.7 0.0132 0.018 0.063210.5 3411.7 0.0144 0.0192 0.064411.5 3736.6 0.0152 0.02 0.065212.5 4061.5 0.0164 0.0212 0.066413.5 4386.4 0.0172 0.022 0.067214.5 4711.3 0.018 0.0228 0.06815.5 5036.3 0.0188 0.0236 0.068816.5 5361.2 0.0196 0.0244 0.069617.5 5686.1 0.0204 0.0252 0.070418.5 6011 0.0212 0.026 0.071219.5 6335.9 0.022 0.0268 0.07220.5 6660.9 0.0228 0.0276 0.072821.5 6985.8 0.0232 0.028 0.073222.5 7310.7 0.024 0.0288 0.07423.5 7635.6 0.0256 0.0304 0.075624.5 7960.5 0.0268 0.0316 0.076825.5 8285.5 0.0288 0.0336 0.078826.5 8610.4 0.03 0.0348 0.0827.5 8935.3 0.0308 0.0356 0.080828.5 9260.2 0.032 0.0368 0.08229.5 9585.1 0.0324 0.0372 0.082430.5 9910.1 0.0328 0.0376 0.082831.5 10235 0.0336 0.0384 0.083632.5 10560 0.034 0.0388 0.08433.5 10885 0.0344 0.0392 0.084434.5 11210 0.0348 0.0396 0.084835.5 11535 0.0356 0.0404 0.085636.5 11860 0.0364 0.0412 0.086437.5 12185 0.0372 0.042 0.087238.5 12509 0.038 0.0428 0.08839.5 12834 0.0388 0.0436 0.088840.5 13159 0.0396 0.0444 0.089641.5 13484 0.0404 0.0452 0.090442.5 13809 0.0412 0.046 0.091244

DSTO-TR-147743.5 14134 0.042 0.0468 0.09244.5 14459 0.0428 0.0476 0.092845.5 14784 0.0436 0.0484 0.093646.5 15109 0.0444 0.0492 0.094447.5 15434 0.0452 0.05 0.095248.5 15759 0.046 0.0508 0.09649.5 16084 0.0468 0.0516 0.096850.5 16408 0.048 0.0528 0.09851.5 16733 0.0488 0.0536 0.098852.5 17058 0.0496 0.0544 0.099653.5 17383 0.0508 0.0556 0.100854.5 17708 0.0516 0.0564 0.101655.5 18033 0.0524 0.0572 0.102456.5 18358 0.0536 0.0584 0.103657.5 18683 0.0544 0.0592 0.104458.5 19008 0.0552 0.06 0.105259.5 19333 0.0564 0.0612 0.106460.5 19658 0.058 0.0628 0.10861.5 19983 0.0596 0.0644 0.109662.5 20307 0.0612 0.066 0.111263.5 20632 0.0624 0.0672 0.112464.5 20957 0.064 0.0688 0.11465.5 21282 0.0656 0.0704 0.115666.5 21607 0.0673 0.0721 0.117367.5 21932 0.0697 0.0745 0.119768.5 22257 0.0713 0.0761 0.121369.5 22582 0.073 0.0778 0.12370.5 22907 0.075 0.0798 0.12571.5 23232 0.0775 0.0823 0.127572.5 23557 0.08 0.0848 0.1373.5 23882 0.0823 0.0871 0.132374.5 24207 0.0841 0.0889 0.134175.5 24531 0.0858 0.0906 0.135876.5 24856 0.0875 0.0923 0.137577.5 25181 0.0895 0.0943 0.139578.5 25506 0.0911 0.0959 0.141179.5 25831 0.0933 0.0981 0.143380.5 26156 0.0956 0.1004 0.145681.5 26481 0.0978 0.1026 0.147882.5 26806 0.0994 0.1042 0.149483.5 27131 0.102 0.1068 0.15284.5 27456 0.1039 0.1087 0.153985.5 27781 0.1066 0.1114 0.156686.5 28106 0.1082 0.113 0.158287.5 28431 0.1093 0.1141 0.159388.5 28755 0.1117 0.1165 0.161789.5 29080 0.1142 0.119 0.164290.5 29405 0.1163 0.1211 0.166391.5 29730 0.1179 0.1227 0.167992.5 30055 0.1198 0.1246 0.169893.5 30380 0.1219 0.1267 0.171994.5 30705 0.1243 0.1291 0.174395.5 31030 0.1278 0.1326 0.177896.5 31355 0.1312 0.136 0.181297.5 31680 0.1326 0.1374 0.182698.5 32005 0.136 0.1408 0.18699.5 32330 0.1391 0.1439 0.1891100.5 32654 0.1392 0.144 0.1892101.5 32979 0.142 0.1468 0.192102.5 33304 0.1435 0.1483 0.1935103.5 33629 0.1452 0.15 0.1952104.5 33954 0.1475 0.1523 0.1975105.5 34279 0.1497 0.1545 0.1997106.5 34604 0.152 0.1568 0.202107.5 34929 0.1547 0.1595 0.2047108.5 35254 0.1566 0.1614 0.206645

DSTO-TR-1477109.5 35579 0.1605 0.1653 0.2105110.5 35904 0.1653 0.1701 0.2153111.5 36229 0.1697 0.1745 0.2197112.5 36553 0.1733 0.1781 0.2233113.5 36878 0.1757 0.1805 0.2257114.5 37203 0.1776 0.1824 0.2276115.5 37528 0.1808 0.1856 0.2308116.5 37853 0.1852 0.19 0.2352117.5 38178 0.1916 0.1964 0.2416118.5 38503 0.1932 0.198 0.2432119.5 38828 0.1981 0.2029 0.2481120.5 39153 0.199 0.2038 0.249121.5 39478 0.2039 0.2087 0.2539122.5 39803 0.2078 0.2126 0.2578123.5 40128 0.2118 0.2166 0.2618124.5 40453 0.2187 0.2235 0.2687125.5 40777 0.2258 0.2306 0.2758126.5 41102 0.2387 0.2435 0.2887127.5 41427 0.2457 0.2505 0.2957128.5 41752 0.2533 0.2581 0.3033129.5 42077 0.2631 0.2679 0.3131130.5 42402 0.2782 0.283 0.3282131.5 42727 0.2943 0.2991 0.3443132.5 43052 0.3163 0.3211 0.3663133.5 43377 0.327 0.3318 0.377134.5 43702 0.3551 0.3599 0.4051135.5 44027 0.3965 0.4013 0.4465136.5 44352 0.4399 0.4447 0.4899137.5 44676 0.4619 0.4667 0.5119138.5 45001 0.5183 0.5231 0.5683139.5 45326 0.5598 0.5646 0.6098140.5 45651 0.6031 0.6079 0.6531141.5 45976 0.6501 0.6549 0.7001142.5 46301 0.7115 0.7163 0.7615143.5 46626 0.7956 0.8004 0.8456144.5 46951 0.9066 0.9114 0.9566145.5 47276 1.0413 1.0461 1.0913146.5 47601 1.2265 1.2313 1.2765147.5 47926 1.4566 1.4614 1.5066148.5 48251 1.9603 1.9651 2.0103149.5 48576 2.801 2.8058 2.851150.5 48900 3.9129 3.9177 3.9629150.74 48978 5.3467 5.3515 5.3967Table A-6. Results of the quantitative fractography on KSIF 140, peak stress 390MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.012 0.032.5 974.76 0.0051 0.0171 0.03513.5 1299.7 0.0058 0.0178 0.03584.5 1624.6 0.0066 0.0186 0.03665.5 1949.5 0.0082 0.0202 0.03826.5 2274.4 0.0093 0.0213 0.03937.5 2599.4 0.0109 0.0229 0.04098.5 2924.3 0.0138 0.0258 0.04389.5 3249.2 0.016 0.028 0.04610.5 3574.1 0.0178 0.0298 0.047811.5 3899 0.0204 0.0324 0.050412.5 4224 0.0258 0.0378 0.055813.5 4548.9 0.0297 0.0417 0.059714.5 4873.8 0.0337 0.0457 0.063715.5 5198.7 0.037 0.049 0.06746

DSTO-TR-147716.5 5523.6 0.0406 0.0526 0.070617.5 5848.6 0.0467 0.0587 0.076718.5 6173.5 0.0493 0.0613 0.079319.5 6498.4 0.0538 0.0658 0.083820.5 6823.3 0.0578 0.0698 0.087821.5 7148.2 0.061 0.073 0.09122.5 7473.2 0.0636 0.0756 0.093623.5 7798.1 0.0687 0.0807 0.098724.5 8123 0.073 0.085 0.10325.5 8447.9 0.0773 0.0893 0.107326.5 8772.8 0.0827 0.0947 0.112727.5 9097.8 0.0863 0.0983 0.116328.5 9422.7 0.0914 0.1034 0.121429.5 9747.6 0.0948 0.1068 0.124830.5 10073 0.0987 0.1107 0.128731.5 10397 0.1057 0.1177 0.135732.5 10722 0.1097 0.1217 0.139733.5 11047 0.1144 0.1264 0.144434.5 11372 0.1207 0.1327 0.150735.5 11697 0.1258 0.1378 0.155836.5 12022 0.1312 0.1432 0.161237.5 12347 0.1367 0.1487 0.166738.5 12672 0.1407 0.1527 0.170739.5 12997 0.1443 0.1563 0.174340.5 13322 0.1502 0.1622 0.180241.5 13647 0.155 0.167 0.18542.5 13972 0.1629 0.1749 0.192943.5 14296 0.1699 0.1819 0.199944.5 14621 0.1761 0.1881 0.206145.5 14946 0.184 0.196 0.21446.5 15271 0.1906 0.2026 0.220647.5 15596 0.1992 0.2112 0.229248.5 15921 0.209 0.221 0.23949.5 16246 0.2168 0.2288 0.246850.5 16571 0.2259 0.2379 0.255951.5 16896 0.2379 0.2499 0.267952.5 17221 0.2507 0.2627 0.280753.5 17546 0.2686 0.2806 0.298654.5 17871 0.2867 0.2987 0.316755.5 18196 0.3111 0.3231 0.341156.5 18520 0.3427 0.3547 0.372757.5 18845 0.3832 0.3952 0.413258.5 19170 0.4333 0.4453 0.463359.5 19495 0.518 0.53 0.54861.5 20145 0.8229 0.8349 0.852963.5 20795 1.1281 1.1401 1.158164.5 21120 1.413 1.425 1.44366.5 21770 2.1305 2.1425 2.160568.5 22419 3.3209 3.3329 3.350969.5 22744 4.3788 4.3908 4.408870.5 22907 5.091 5.103 5.121Table A-7. Results of the quantitative fractography on KSIF 205, peak stress 390MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0324 0.071.5 487.38 0.002 0.0344 0.0722.5 812.3 0.004 0.0364 0.0743.5 1137.2 0.0068 0.0392 0.07684.5 1462.1 0.0092 0.0416 0.07925.5 1787.1 0.012 0.0444 0.0826.5 2112 0.0168 0.0492 0.086847

DSTO-TR-14777.5 2436.9 0.0204 0.0528 0.09048.5 2761.8 0.0232 0.0556 0.09329.5 3086.7 0.0288 0.0612 0.098810.5 3411.7 0.0356 0.068 0.105611.5 3736.6 0.0388 0.0712 0.108812.5 4061.5 0.046 0.0784 0.11613.5 4386.4 0.0516 0.084 0.121614.5 4711.3 0.0564 0.0888 0.126415.5 5036.3 0.0628 0.0952 0.132816.5 5361.2 0.0672 0.0996 0.137217.5 5686.1 0.0722 0.1046 0.142218.5 6011 0.0786 0.111 0.148619.5 6335.9 0.0822 0.1146 0.152220.5 6660.9 0.0882 0.1206 0.158221.5 6985.8 0.0926 0.125 0.162622.5 7310.7 0.0986 0.131 0.168623.5 7635.6 0.1042 0.1366 0.174224.5 7960.5 0.1094 0.1418 0.179425.5 8285.5 0.1142 0.1466 0.184226.5 8610.4 0.1178 0.1502 0.187827.5 8935.3 0.1225 0.1549 0.192528.5 9260.2 0.1289 0.1613 0.198929.5 9585.1 0.1365 0.1689 0.206530.5 9910.1 0.1433 0.1757 0.213331.5 10235 0.1537 0.1861 0.223732.5 10560 0.1617 0.1941 0.231733.5 10885 0.1713 0.2037 0.241334.5 11210 0.1833 0.2157 0.253335.5 11535 0.1957 0.2281 0.265736.5 11860 0.2109 0.2433 0.280937.5 12185 0.2261 0.2585 0.296138.5 12509 0.2457 0.2781 0.315739.5 12834 0.2661 0.2985 0.336140.5 13159 0.2893 0.3217 0.359341.5 13484 0.3125 0.3449 0.382542.5 13809 0.3445 0.3769 0.414543.5 14134 0.3852 0.4176 0.455244.5 14459 0.4608 0.4932 0.530845.5 14784 0.5643 0.5967 0.634346.5 15109 0.668 0.7004 0.73847.5 15434 0.8125 0.8449 0.882548.5 15759 1.0824 1.1148 1.152449.5 16084 1.4658 1.4982 1.535850.5 16408 2.0536 2.086 2.123651.5 16733 2.9335 2.9659 3.003552.2 16961 3.5692 3.6016 3.6392Table A-8. Results of the quantitative fractography on KSIF 118, peak stress 390MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0188 0.020.5 162.46 0.0016 0.0204 0.02162.5 812.3 0.004 0.0228 0.0244.5 1462.1 0.0056 0.0244 0.02565.5 1787.1 0.0072 0.026 0.02726.5 2112 0.0088 0.0276 0.02888.5 2761.8 0.0116 0.0304 0.031610.5 3411.7 0.014 0.0328 0.03411.5 3736.6 0.0156 0.0344 0.035612.5 4061.5 0.0172 0.036 0.037214.5 4711.3 0.0204 0.0392 0.040416.5 5361.2 0.0232 0.042 0.043248

DSTO-TR-147718.5 6011 0.0264 0.0452 0.046420.5 6660.9 0.0292 0.048 0.049221.5 6985.8 0.032 0.0508 0.05222.5 7310.7 0.0344 0.0532 0.054424.5 7960.5 0.0376 0.0564 0.057626.5 8610.4 0.0408 0.0596 0.060828.5 9260.2 0.0476 0.0664 0.067629.5 9585.1 0.0512 0.07 0.071234.5 11210 0.0698 0.0886 0.089836.5 11860 0.0752 0.094 0.095237.5 12185 0.0796 0.0984 0.099638.5 12509 0.0852 0.104 0.105239.5 12834 0.0908 0.1096 0.110840.5 13159 0.0972 0.116 0.117241.5 13484 0.1008 0.1196 0.120842.5 13809 0.106 0.1248 0.12643.5 14134 0.1104 0.1292 0.130444.5 14459 0.1176 0.1364 0.137645.5 14784 0.1228 0.1416 0.142846.5 15109 0.1264 0.1452 0.146447.5 15434 0.1308 0.1496 0.150848.5 15759 0.1348 0.1536 0.154849.5 16084 0.1396 0.1584 0.159650.5 16408 0.146 0.1648 0.16651.5 16733 0.152 0.1708 0.17252.5 17058 0.16 0.1788 0.1853.5 17383 0.166 0.1848 0.18654.5 17708 0.1728 0.1916 0.192855.5 18033 0.1792 0.198 0.199256.5 18358 0.1888 0.2076 0.208857.5 18683 0.1972 0.216 0.217258.5 19008 0.2064 0.2252 0.226459.5 19333 0.2168 0.2356 0.236860.5 19658 0.2344 0.2532 0.254461.5 19983 0.2464 0.2652 0.266462.5 20307 0.2652 0.284 0.285263.5 20632 0.2848 0.3036 0.304864.5 20957 0.308 0.3268 0.32865.5 21282 0.3326 0.3514 0.352666.5 21607 0.355 0.3738 0.37567.5 21932 0.3794 0.3982 0.399468.5 22257 0.4034 0.4222 0.423469.5 22582 0.429 0.4478 0.44970.5 22907 0.4628 0.4816 0.482871.5 23232 0.507 0.5258 0.52772.5 23557 0.5883 0.6071 0.608373.5 23882 0.7708 0.7896 0.790874.5 24207 0.9631 0.9819 0.983175.5 24531 1.1784 1.1972 1.198476.5 24856 1.4513 1.4701 1.471377.5 25181 1.861 1.8798 1.88178.5 25506 2.5415 2.5603 2.561579.5 25831 3.1274 3.1462 3.147480.5 26156 3.5793 3.5981 3.5993Table A-9. Results of the quantitative fractography on KSIF 116, peak stress 390MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.036 0.0450.5 162.46 0.0056 0.0416 0.05061.5 487.38 0.0076 0.0436 0.05262.5 812.3 0.0092 0.0452 0.054249

DSTO-TR-14773.5 1137.2 0.0116 0.0476 0.05664.5 1462.1 0.0132 0.0492 0.05825.5 1787.1 0.0156 0.0516 0.06066.5 2112 0.0188 0.0548 0.06387.5 2436.9 0.0216 0.0576 0.06668.5 2761.8 0.0244 0.0604 0.06949.5 3086.7 0.026 0.062 0.07110.5 3411.7 0.0284 0.0644 0.073411.5 3736.6 0.0304 0.0664 0.075412.5 4061.5 0.0328 0.0688 0.077813.5 4386.4 0.0358 0.0718 0.080814.5 4711.3 0.0377 0.0737 0.082717.5 5686.1 0.0467 0.0827 0.091730.5 9910.1 0.0932 0.1292 0.138233.5 10885 0.1074 0.1434 0.152434.5 11210 0.1122 0.1482 0.157235.5 11535 0.1178 0.1538 0.162836.5 11860 0.1253 0.1613 0.170337.5 12185 0.1305 0.1665 0.175538.5 12509 0.1381 0.1741 0.183139.5 12834 0.1497 0.1857 0.194740.5 13159 0.1625 0.1985 0.207541.5 13484 0.1721 0.2081 0.217142.5 13809 0.1801 0.2161 0.225143.5 14134 0.1865 0.2225 0.231544.5 14459 0.1933 0.2293 0.238345.5 14784 0.2077 0.2437 0.252746.5 15109 0.2285 0.2645 0.273547.5 15434 0.2485 0.2845 0.293548.5 15759 0.2705 0.3065 0.315549.5 16084 0.2905 0.3265 0.335550.5 16408 0.3129 0.3489 0.357951.5 16733 0.3369 0.3729 0.381952.5 17058 0.3729 0.4089 0.417953.5 17383 0.4176 0.4536 0.462654.5 17708 0.4615 0.4975 0.506555.5 18033 0.536 0.572 0.58156.5 18358 0.6232 0.6592 0.668257.5 18683 0.7582 0.7942 0.803258.5 19008 0.919 0.955 0.96459.5 19333 1.1831 1.2191 1.228160.5 19658 1.6982 1.7342 1.743261.5 19983 2.2517 2.2877 2.296762.5 20307 3.0663 3.1023 3.111363.05 20486 4.2993 4.3353 4.3443Table A-10. Results of the quantitative fractography on KSIF 151, peak stress 390MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0368 0.0310.5 3411.7 0.0142 0.051 0.044211.5 3736.6 0.0162 0.053 0.046213.5 4386.4 0.0202 0.057 0.050214.5 4711.3 0.0218 0.0586 0.051815.5 5036.3 0.0238 0.0606 0.053816.5 5361.2 0.0262 0.063 0.056217.5 5686.1 0.0281 0.0649 0.058118.5 6011 0.0317 0.0685 0.061719.5 6335.9 0.0341 0.0709 0.064120.5 6660.9 0.0381 0.0749 0.068121.5 6985.8 0.0405 0.0773 0.070522.5 7310.7 0.0441 0.0809 0.074150

DSTO-TR-147723.5 7635.6 0.0481 0.0849 0.078124.5 7960.5 0.051 0.0878 0.08125.5 8285.5 0.0534 0.0902 0.083426.5 8610.4 0.0564 0.0932 0.086427.5 8935.3 0.0611 0.0979 0.091128.5 9260.2 0.0647 0.1015 0.094729.5 9585.1 0.0671 0.1039 0.097130.5 9910.1 0.0711 0.1079 0.101131.5 10235 0.0751 0.1119 0.105132.5 10560 0.0799 0.1167 0.109933.5 10885 0.0846 0.1214 0.114634.5 11210 0.0886 0.1254 0.118635.5 11535 0.0946 0.1314 0.124636.5 11860 0.1006 0.1374 0.130637.5 12185 0.1074 0.1442 0.137438.5 12509 0.113 0.1498 0.14339.5 12834 0.1202 0.157 0.150240.5 13159 0.1298 0.1666 0.159841.5 13484 0.1373 0.1741 0.167342.5 13809 0.1461 0.1829 0.176143.5 14134 0.1553 0.1921 0.185344.5 14459 0.1657 0.2025 0.195745.5 14784 0.1757 0.2125 0.205746.5 15109 0.1901 0.2269 0.220147.5 15434 0.2061 0.2429 0.236148.5 15759 0.2253 0.2621 0.255349.5 16084 0.2453 0.2821 0.275350.5 16408 0.2693 0.3061 0.299351.5 16733 0.3005 0.3373 0.330552.5 17058 0.3335 0.3703 0.363553.5 17383 0.3669 0.4037 0.396954.5 17708 0.4084 0.4452 0.438455.5 18033 0.4669 0.5037 0.496956.5 18358 0.5555 0.5923 0.585557.5 18683 0.6627 0.6995 0.692758.5 19008 0.8245 0.8613 0.854559.5 19333 1.0262 1.063 1.056260.5 19658 1.291 1.3278 1.32161.5 19983 1.5902 1.627 1.620262.5 20307 2.1388 2.1756 2.168863.5 20632 2.4998 2.5366 2.529864.5 20957 3.2487 3.2855 3.278765.5 21282 4.0363 4.0731 4.0663Table A-11. Results of the quantitative fractography on KSIF 142, peak stress 420MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.056 0.050.5 162.46 0.002 0.058 0.0521.5 487.38 0.0068 0.0628 0.05682.5 812.3 0.0116 0.0676 0.06163.5 1137.2 0.0184 0.0744 0.06844.5 1462.1 0.0264 0.0824 0.07645.5 1787.1 0.034 0.09 0.0846.5 2112 0.0444 0.1004 0.09447.5 2436.9 0.0532 0.1092 0.10328.5 2761.8 0.0652 0.1212 0.11529.5 3086.7 0.0772 0.1332 0.127210.5 3411.7 0.09 0.146 0.1411.5 3736.6 0.106 0.162 0.15612.5 4061.5 0.1235 0.1795 0.173513.5 4386.4 0.1387 0.1947 0.188751

DSTO-TR-147714.5 4711.3 0.1565 0.2125 0.206515.5 5036.3 0.1736 0.2296 0.223616.5 5361.2 0.1955 0.2515 0.245517.5 5686.1 0.2254 0.2814 0.275418.5 6011 0.2568 0.3128 0.306819.5 6335.9 0.2981 0.3541 0.348120.5 6660.9 0.3496 0.4056 0.399621.5 6985.8 0.4197 0.4757 0.469722.5 7310.7 0.5158 0.5718 0.565823.5 7635.6 0.6494 0.7054 0.699424.5 7960.5 0.9213 0.9773 0.971325.5 8285.5 1.3358 1.3918 1.385826.5 8610.4 2.0847 2.1407 2.134728.4 9227.7 4.2004 4.2564 4.2504Table A-12. Results of the quantitative fractography on KSIF 204, peak stress 420MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0168 0.030.5 162.46 0.002 0.0188 0.0321.5 487.38 0.006 0.0228 0.0362.5 812.3 0.0112 0.028 0.04123.5 1137.2 0.018 0.0348 0.0484.5 1462.1 0.0248 0.0416 0.05485.5 1787.1 0.0296 0.0464 0.05966.5 2112 0.0338 0.0506 0.06387.5 2436.9 0.0398 0.0566 0.06988.5 2761.8 0.0492 0.066 0.079210.5 3411.7 0.0626 0.0794 0.092611.5 3736.6 0.0698 0.0866 0.099812.5 4061.5 0.0801 0.0969 0.110114.5 4711.3 0.0947 0.1115 0.124715.5 5036.3 0.1018 0.1186 0.131816.5 5361.2 0.1084 0.1252 0.138417.5 5686.1 0.1187 0.1355 0.148718.5 6011 0.1278 0.1446 0.157819.5 6335.9 0.1377 0.1545 0.167720.5 6660.9 0.1471 0.1639 0.177121.5 6985.8 0.1633 0.1801 0.193322.5 7310.7 0.1833 0.2001 0.213323.5 7635.6 0.2101 0.2269 0.240124.5 7960.5 0.2455 0.2623 0.275525.5 8285.5 0.2886 0.3054 0.318626.5 8610.4 0.3333 0.3501 0.363327.5 8935.3 0.4123 0.4291 0.442328.5 9260.2 0.5004 0.5172 0.530429.5 9585.1 0.6455 0.6623 0.675530.5 9910.1 0.8154 0.8322 0.845431.5 10235 1.0712 1.088 1.101232.5 10560 1.5158 1.5326 1.545833.5 10885 2.0603 2.0771 2.090334.5 11210 2.815 2.8318 2.84535.13 11414 3.6563 3.6731 3.686352

DSTO-TR-1477Table A-13. Results of the quantitative fractography on KSIF 147, peak stress 420MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0308 0.031.5 487.38 0.0029 0.0337 0.03293.5 1137.2 0.0079 0.0387 0.03794.5 1462.1 0.0118 0.0426 0.04185.5 1787.1 0.0162 0.047 0.04626.5 2112 0.0197 0.0505 0.04977.5 2436.9 0.0234 0.0542 0.05348.5 2761.8 0.0288 0.0596 0.05889.5 3086.7 0.0335 0.0643 0.063510.5 3411.7 0.0373 0.0681 0.067311.5 3736.6 0.0403 0.0711 0.070312.5 4061.5 0.0448 0.0756 0.074813.5 4386.4 0.049 0.0798 0.07914.5 4711.3 0.054 0.0848 0.08415.5 5036.3 0.0606 0.0914 0.090616.5 5361.2 0.0676 0.0984 0.097617.5 5686.1 0.0767 0.1075 0.106718.5 6011 0.0873 0.1181 0.117319.5 6335.9 0.0982 0.129 0.128220.5 6660.9 0.1118 0.1426 0.141821.5 6985.8 0.1301 0.1609 0.160122.5 7310.7 0.1484 0.1792 0.178423.5 7635.6 0.1638 0.1946 0.193824.5 7960.5 0.1896 0.2204 0.219626.5 8610.4 0.2471 0.2779 0.277127.5 8935.3 0.2877 0.3185 0.317728.5 9260.2 0.3392 0.37 0.369229.5 9585.1 0.4031 0.4339 0.433130.5 9910.1 0.5078 0.5386 0.537831.5 10235 0.6324 0.6632 0.662432.5 10560 0.797 0.8278 0.82733.5 10885 0.97 1.0008 134.5 11210 1.3165 1.3473 1.346535.5 11535 1.7938 1.8246 1.823836.5 11860 2.4968 2.5276 2.526837.3 12120 3.9253 3.9561 3.9553Table A-14. Results of the quantitative fractography on KSIF 193, peak stress 420MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0284 0.030.5 162.46 0.0044 0.0328 0.03441.5 487.38 0.008 0.0364 0.0382.5 812.3 0.01 0.0384 0.043.5 1137.2 0.0156 0.044 0.04564.5 1462.1 0.0184 0.0468 0.04845.5 1787.1 0.0228 0.0512 0.05286.5 2112 0.0272 0.0556 0.05727.5 2436.9 0.0316 0.06 0.06168.5 2761.8 0.036 0.0644 0.0669.5 3086.7 0.0408 0.0692 0.070810.5 3411.7 0.0456 0.074 0.075611.5 3736.6 0.0512 0.0796 0.081212.5 4061.5 0.0576 0.086 0.087613.5 4386.4 0.0644 0.0928 0.094453

DSTO-TR-147714.5 4711.3 0.0708 0.0992 0.100815.5 5036.3 0.0776 0.106 0.107616.5 5361.2 0.0856 0.114 0.115617.5 5686.1 0.0952 0.1236 0.125218.5 6011 0.1056 0.134 0.135619.5 6335.9 0.1184 0.1468 0.148420.5 6660.9 0.1332 0.1616 0.163221.5 6985.8 0.1488 0.1772 0.178822.5 7310.7 0.1684 0.1968 0.198423.5 7635.6 0.192 0.2204 0.22224.5 7960.5 0.2192 0.2476 0.249225.5 8285.5 0.2488 0.2772 0.278826.5 8610.4 0.2844 0.3128 0.314427.5 8935.3 0.3252 0.3536 0.355228.5 9260.2 0.3802 0.4086 0.410229.5 9585.1 0.4517 0.4801 0.481730.5 9910.1 0.526 0.5544 0.55631.5 10235 0.6367 0.6651 0.666732.5 10560 0.8363 0.8647 0.866333.5 10885 1.061 1.0894 1.09134.5 11210 1.2599 1.2883 1.289936.5 11860 2.5209 2.5493 2.550937.4 12152 3.1174 3.1458 3.1474Table A-15. Results of the quantitative fractography on KSIF 191, peak stress 420MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0208 0.040.5 162.46 0.0044 0.0252 0.04441.5 487.38 0.0108 0.0316 0.05082.5 812.3 0.018 0.0388 0.0583.5 1137.2 0.0232 0.044 0.06324.5 1462.1 0.0292 0.05 0.06925.5 1787.1 0.036 0.0568 0.0766.5 2112 0.042 0.0628 0.0827.5 2436.9 0.0464 0.0672 0.08648.5 2761.8 0.0536 0.0744 0.09369.5 3086.7 0.0604 0.0812 0.100410.5 3411.7 0.0688 0.0896 0.108811.5 3736.6 0.0804 0.1012 0.120412.5 4061.5 0.0888 0.1096 0.128813.5 4386.4 0.098 0.1188 0.13814.5 4711.3 0.1092 0.13 0.149215.5 5036.3 0.12 0.1408 0.1616.5 5361.2 0.1328 0.1536 0.172817.5 5686.1 0.1476 0.1684 0.187618.5 6011 0.1644 0.1852 0.204419.5 6335.9 0.1844 0.2052 0.224420.5 6660.9 0.208 0.2288 0.24821.5 6985.8 0.2364 0.2572 0.276422.5 7310.7 0.28 0.3008 0.3223.5 7635.6 0.322 0.3428 0.36224.5 7960.5 0.3664 0.3872 0.406425.5 8285.5 0.4393 0.4601 0.479326.5 8610.4 0.5191 0.5399 0.559127.5 8935.3 0.6491 0.6699 0.689128.5 9260.2 0.8331 0.8539 0.873129.5 9585.1 1.0992 1.12 1.139230.5 9910.1 1.5514 1.5722 1.591431.5 10235 2.3776 2.3984 2.417632.1 10430 3.0873 3.1081 3.127354

DSTO-TR-1477Table A-16. Results of the quantitative fractography on KSIF 183, peak stress 450MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.028 0.030.5 162.46 0.006 0.034 0.0362.5 812.3 0.016 0.044 0.0463.5 1137.2 0.0244 0.0524 0.05444.5 1462.1 0.0348 0.0628 0.06485.5 1787.1 0.0433 0.0713 0.07336.5 2112 0.0561 0.0841 0.08617.5 2436.9 0.0684 0.0964 0.09848.5 2761.8 0.0839 0.1119 0.11399.5 3086.7 0.1003 0.1283 0.130310.5 3411.7 0.1222 0.1502 0.152211.5 3736.6 0.1562 0.1842 0.186212.5 4061.5 0.2014 0.2294 0.231413.5 4386.4 0.2462 0.2742 0.276214.5 4711.3 0.2933 0.3213 0.323315.5 5036.3 0.3973 0.4253 0.427316.5 5361.2 0.58 0.608 0.6118.5 6011 1.272 1.3 1.30219.4 6303.4 1.948 1.976 1.978Table A-17. Results of the quantitative fractography on KSIF 178, peak stress 450MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.013 0.011.5 487.38 0.0116 0.0246 0.02162.5 812.3 0.0204 0.0334 0.03043.5 1137.2 0.0348 0.0478 0.04484.5 1462.1 0.0597 0.0727 0.06975.5 1787.1 0.0963 0.1093 0.10636.5 2112 0.1433 0.1563 0.15337.5 2436.9 0.1997 0.2127 0.20978.5 2761.8 0.2565 0.2695 0.26659.5 3086.7 0.3583 0.3713 0.368310.5 3411.7 0.5123 0.5253 0.522311.5 3736.6 0.6954 0.7084 0.705412.5 4061.5 0.8753 0.8883 0.885313.5 4386.4 1.3675 1.3805 1.377514.73 4786.1 1.8905 1.9035 1.9005Table A-18. Results of the quantitative fractography on KSIF 186, peak stress 450MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0392 0.030.5 162.46 0.004 0.0432 0.0341.5 487.38 0.008 0.0472 0.0382.5 812.3 0.0132 0.0524 0.04323.5 1137.2 0.0192 0.0584 0.04924.5 1462.1 0.0252 0.0644 0.05525.5 1787.1 0.034 0.0732 0.0646.5 2112 0.0464 0.0856 0.07647.5 2436.9 0.0608 0.1 0.09088.5 2761.8 0.0836 0.1228 0.113655

DSTO-TR-14779.5 3086.7 0.1092 0.1484 0.139210.5 3411.7 0.144 0.1832 0.17411.5 3736.6 0.1818 0.221 0.211812.5 4061.5 0.2508 0.29 0.280813.5 4386.4 0.3673 0.4065 0.397314.5 4711.3 0.5248 0.564 0.554815.5 5036.3 0.7532 0.7924 0.783216.5 5361.2 1.1665 1.2057 1.196517.5 5686.1 1.9283 1.9675 1.958319.3 6271 3.3166 3.3558 3.3466Table A-19. Results of the quantitative fractography on KSIF 164, peak stress 450MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0076 0.015.5 1787.1 0.029 0.0366 0.0396.5 2112 0.0347 0.0423 0.04477.5 2436.9 0.0406 0.0482 0.05069.5 3086.7 0.0578 0.0654 0.067810.5 3411.7 0.0766 0.0842 0.086611.5 3736.6 0.0962 0.1038 0.106212.5 4061.5 0.1254 0.133 0.135413.5 4386.4 0.1486 0.1562 0.158614.5 4711.3 0.1882 0.1958 0.198215.5 5036.3 0.2558 0.2634 0.265816.5 5361.2 0.3426 0.3502 0.352617.5 5686.1 0.5006 0.5082 0.510618.5 6011 0.757 0.7646 0.76719.5 6335.9 1.3291 1.3367 1.339120.5 6660.9 1.9167 1.9243 1.926720.9 6790.8 2.424 2.4316 2.434Table A-20. Results of the quantitative fractography on KSIF 146, peak stress 450MPaNo. of spectraIARPO3a + 5 markerloadsSimulated FlightHoursMeasurement fromflaw edge to crackpositionmm.Addition of Type IEPS (measured flawdepth)mmAddition of Type IIEPSmm0 0 0.0248 0.050.5 162.46 0.0072 0.032 0.05721.5 487.38 0.0148 0.0396 0.06482.5 812.3 0.0228 0.0476 0.07283.5 1137.2 0.0352 0.06 0.08524.5 1462.1 0.044 0.0688 0.0945.5 1787.1 0.0548 0.0796 0.10486.5 2112 0.0636 0.0884 0.11367.5 2436.9 0.0756 0.1004 0.12568.5 2761.8 0.0944 0.1192 0.14449.5 3086.7 0.1137 0.1385 0.163710.5 3411.7 0.1406 0.1654 0.190611.5 3736.6 0.1743 0.1991 0.224312.5 4061.5 0.2109 0.2357 0.260913.5 4386.4 0.2635 0.2883 0.313514.5 4711.3 0.3352 0.36 0.385215.5 5036.3 0.4479 0.4727 0.497917.5 5686.1 0.7486 0.7734 0.798618.5 6011 1.0677 1.0925 1.117719.5 6335.9 1.4946 1.5194 1.544620 6498.4 2.6592 2.684 2.709256

DISTRIBUTION LISTFatigue Crack Growth in 7050T7451 Aluminium Alloy Thick Section Plate with a Glass BeadPeened Surface Simulating Some Regions of the F/A-18 StructureAUSTRALIADEFENCE ORGANISATIONS A BarterTask Sponsor DGTA-ASI1 (SQNLDR J. Medved) 2S&T ProgramChief Defence Scientist⎫FAS Science Policy ⎬ shared copy 1AS Science Corporate Management ⎭Director General Science Policy Development 1Counsellor Defence Science, London (Doc Data Sheet )Counsellor Defence Science, Washington (Doc Data Sheet )Scientific Adviser to MRDC Thailand (Doc Data Sheet )Scientific Adviser Policy and Command 1Navy Scientific Adviser (Doc Data Sheet and distribution list only)Scientific Adviser - Army (Doc Data Sheet and distribution list only)Air Force Scientific Adviser 1Director Trials 1Platforms Sciences LaboratoryDirector 1Chief of AVD. Division 1Research Leader AVD-ASI 1Task Manager L Molent 1Author: S A Barter 1P White 1K Walker 1Q Liu 1P K Sharp 1IFOSTP filing system 1DSTO Library and ArchivesLibrary Fishermans Bend 1Library Maribyrnong 1Library Salisbury 2Australian Archives 1Library, MOD, Pyrmont (Doc Data sheet only)Air ForceTFSPO Williamtown (Chief Engineer) 1RAAF TLO IFOSTP (SQNLDR P. Klose) 1

Intelligence ProgramDGSTA Defence Intelligence Organisation 1Manager, Information Centre, Defence Intelligence Organisation 1CanadaDTA 1NRC – D Simpson 1Bombardier – Y Richard 1SPARES 5Total number of copies: 34

Page classification: UNCLASSIFIEDDEFENCE SCIENCE AND TECHNOLOGY ORGANISATIONDOCUMENT CONTROL DATA1. PRIVACY MARKING/CAVEAT (OF DOCUMENT)2. TITLEFatigue Crack Growth in 7050T7451 Aluminium Alloy Thick SectionPlate with a Glass Bead Peened Surface Simulating Some Regions ofthe F/A-18 Structure4. AUTHOR(S)S. A. Barter3. SECURITY CLASSIFICATION (FOR UNCLASSIFIED REPORTSTHAT ARE LIMITED RELEASE USE (L) NEXT TO DOCUMENTCLASSIFICATION)DocumentTitleAbstract5. CORPORATE AUTHOR(U)(U)(U)Platforms Sciences Laboratory506 Lorimer StFishermans Bend Victoria 3207 Australia6a. DSTO NUMBERDSTO-TR-14776b. AR NUMBERAR-012-8566c. TYPE OF REPORTTechnical Report7. DOCUMENT DATEAugust 20038. FILE NUMBER2003/36886/19. TASK NUMBERAIR0014110. TASK SPONSORASI11. NO. OF PAGES5612. NO. OF REFERENCES2213. URL on the World Wide Web14. RELEASE AUTHORITYhttp://www.dsto.defence.gov.au/corporate/reports/DSTO-TR-1477.pdfChief, Air Vehicles Division15. SECONDARY RELEASE STATEMENT OF THIS DOCUMENTApproved for public releaseOVERSEAS ENQUIRIES OUTSIDE STATED LIMITATIONS SHOULD BE REFERRED THROUGH DOCUMENT EXCHANGE, PO BOX 1500, EDINBURGH, SA 511116. DELIBERATE ANNOUNCEMENTNo Limitations17. CITATION IN OTHER DOCUMENTS Yes18. DEFTEST DESCRIPTORSF/A-18 Aircraft, Aluminium Alloy, Shot Peening, Cracking, Fatigue19. ABSTRACTThis report presents the results of a fatigue coupon test program whose primary purpose was to obtain resultsfrom coupons treated with a glass bead peened surfaces typical of some regions of critical F/A-18 aircraftstructure. A spectrum representative of RAAF’s F/A-18 fleet fatigue usage was used. The coupons wererepresentative of the material and geometry of a structural detail that has been found to be fatigue-critical inthe 7050T7451 high strength aluminium alloy wing carry through bulkheads. Following the tests, quantitativefractography was used to produce crack growth curves for each of the fatigued specimens. This reportdescribes the surface condition being examined, test spectrum, test methods, test results and examines ways ofinterpreting the crack growth curves to establish a measure of the severity of the flaws from which the fatiguecracks initiated.Page classification: UNCLASSIFIED