Failure Analysis

National Aerospace Laboratory NLR (Netherlands) and Defence Science and Technology Organisation DSTO (Australia)

Fatigue failure analyses for metallic aircraft comp onents

Russell Wanhill, Simon Barter, Lorrie Molent

National Aerospace Laboratory NLR (Netherlands) and Defence Science and Technology Organisation DSTO (Australia) 19-3-2014 2

Introduction: failure modes and frequencies


Failure modes in aircraft

Failure modes Percentages of failuresCorrosionFatigueBrittle fractureOverload (fast fracture)High temperature corrosionSCC/CF/HECreepWear/abrasion/erosion

3-1655-61

–14-18

27-81

6-7

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* SCC: stress corrosion cracking; CF: corrosion fatig ue; HE: hydrogen embrittlement

Sources: C.R. Brooks and A. Choudhury (2001); S.J. Findley and N.D. Harrison (2002)


Threats to aircraft structural safety

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Threats Failure typesHigh local stresses

Fabrication/material defects

Maintenance damage/deficiencies

Widespread Fatigue Damage (ageing aircraft)

Environmental damage

High operational loads

Impact (various sources)

Explosive/ballistic penetrations

fatigue

fatigue

fatigue

fatigue

corrosion, stress corrosion

overload

abrupt cracking, fatigue

abrupt cracking

Source: C.F. Tiffany, J.P. Gallagher and C.A. Babish,IV (2010)


Failure modes in aircraft: summary

� Fatigue failures predominate in aircraft structures and ar e alsothe main contributors to structural accidents and incident s*

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*Sources: G.S. Campbell and A.R. Lahey (1984); R.J.H. Wanhill (2009);C.F. Tiffany, J.P. Gallagher and C.A. Babish, IV (2010)


Fatigue “initiation”


Overview of fatigue-initiating discontinuities

� Aircraft structures and components have many sources ofdiscontinuities capable of causing fatigue cracking very q uickly *

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Discontinuity sources Specifics

Poorly finished holes

poor drillingscoring from fastenerspoor de-burringmachining tears/nicks

Surface treatments

etch pits; etch pits + machiningintergranular ‘penetration’pickling; chemical milling + peeningpeening laps and cuts

Constituent particles cracked particlesvarying shapesparticles + poor machining

Porosity thick plate mid-line porosity

*Source: S.A. Barter, L. Molent and R.J.H. Wanhill (2012)


Fatigue crack growth (FCG)


FCG characteristics in aircraft structures and comp onents *:Quantitative Fractography (QF) results

� FCG in highly stressed locations and areas often oc curs▬ from small discontinuities ≤ 0.1mm: ‘short cracks’ to ‘long’ cracks▬ almost immediately upon entering service

� ‘Lead cracks’ often show approximately exponential FCGirrespective of the fatigue load history, aircraft type andstructural geometry: see slide 10 for examples

� Environmental effects (corrosion fatigue) are unusu al or minimal compared to fatigue in normal air

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*Sources: L. Molent, S.A. Barter and R.J.H. Wanhill (2010, 2011)


Approximately exponential FCG at many locationsin the lower wing skin of a General Dynamics F-111

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0.0001

0.001

0.01

0.1

1

10

0 5000 10000 15000 20000 25000 30000 35000 40000Service + Test Hours

Cra

ck d

epth

(m

m)

IPP (1) IPP (2)IPP (3) IPP (4)

IPP (5) IPP (6)IPP (6_A) FASS 226 (2)

FASS 226 (4) FASS 226 (5)FASS 226 (5L) FASS 226 (6)

FASS 226 (6R) SIH_OBD_SPLICESIH_FS238/239 (1) SIH_FS238/239 (2)

SIH_FS238/239 (3) SIH_FS238/239 (4)SIH_FS238/239 (5) SPLICE 230 (Rad)

SPLICE 230 (Skin) SPLICE 230 (Bore)SPLICE 244 AAS084 (1)

AAS084 (2) CS086 (1c)CS086 (1d) CS086 (2)

CS100 (1) CS100 (2)DI 95 - BLKHD 1 FWD Hole CS093 (skin)

SPLICE 228 SIH_FAS173SPLICE 239 SIH FS242B

SPLICE 231 SPLICE 237SIH FS230_2 SIH FS230_1

FASS 281 (1) FASS 281 (2)FASS 281 (6) FASS 281 (7)

FS263A RS137ACS043 (1,A15-5) CS043 (8,A15-5)BLKHD 1 AFT HOLE

TestService

0.0001

0.001

0.01

0.1

1

10

0 5000 10000 15000 20000 25000 30000 35000 40000Service + Test Hours

Cra

ck d

epth

(m

m)

IPP (1) IPP (2)IPP (3) IPP (4)

IPP (5) IPP (6)IPP (6_A) FASS 226 (2)

FASS 226 (4) FASS 226 (5)FASS 226 (5L) FASS 226 (6)

FASS 226 (6R) SIH_OBD_SPLICESIH_FS238/239 (1) SIH_FS238/239 (2)

SIH_FS238/239 (3) SIH_FS238/239 (4)SIH_FS238/239 (5) SPLICE 230 (Rad)

SPLICE 230 (Skin) SPLICE 230 (Bore)SPLICE 244 AAS084 (1)

AAS084 (2) CS086 (1c)CS086 (1d) CS086 (2)

CS100 (1) CS100 (2)DI 95 - BLKHD 1 FWD Hole CS093 (skin)

SPLICE 228 SIH_FAS173SPLICE 239 SIH FS242B

SPLICE 231 SPLICE 237SIH FS230_2 SIH FS230_1

FASS 281 (1) FASS 281 (2)FASS 281 (6) FASS 281 (7)

FS263A RS137ACS043 (1,A15-5) CS043 (8,A15-5)BLKHD 1 AFT HOLE

TestServiceService

Service + Test Hours

Test

Cra

ck d

epth

(m

m)

40,00035,00030,00025,00020,00015,00010,0005000

0.001

0.01

0.1

10

1

0.00010


Potential and actual uses of FCG analyses - I

� Service failures

� Full-scale and component tests: ▬ interim -detected cracking▬ teardown-detected cracks

● Special cases, e.g.▬ proof testing of F-111s in service:

Buntin (1977), Redmond (2001)

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F/A-18 centre barrel bulkhead

Note: aircraft FCG often occurs instructures and components havingcomplex geometries and/or loads: seethe next slide also


Aloha Airlines Boeing 737 accident owingto Multiple Site fatigue Damage (MSD)

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Potential and actual uses of FCG analyses - II

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● Life to maximum / critical cracksize, ac , from :▬ initial discontinuity/crack size, ai

▬ NDI crack size limit, andi

● Feasibility of repair / replacementfor remaining service or test life

● Effect of aircraft usage severity(individual and fleet differences)▬ continued safe operation


FCG analysis methods


Types of FCG analyses

� LEFM based analyses

� Non-LEFM based analyses

� LEFM / non-LEFM based analyses

▬ ad hoc ‘hybrid’ analyses

▬ Effective Block Approach ( EBA )

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Survey of DSTO and NLR FCG analysis examples

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FCG analysis type Failure/cracking category DSTO NLRLEFM

Non-LEFM

LEFM /non-LEFM▬ ‘hybrid’ analysis▬ EBA

service

servicetest

servicetest only

–

46

1

–

37

1–

Non-LEFM FCG analysis types

DSTO

NLR

analytical: based on a versus N plots and

numerical

analytical: based on da/dN versus a plots and

Nf 0a a eλ=

BadaAe

dN=

� Exclusively LEFM based analyses were /are not used: see the next slide


Selection and use of FCG analysis methods(Wanhill, Molent and Barter 2012/13)

� Fracture mechanics (LEFM based) analyses are▬ unsuitable for direct analysis of service failure FCG ▬ sometimes useful, when calibrated against QF a versus N test data, for

estimates of total FCG lives to failure and effect of usage severity▬ useful, even essential, for estimates of maximum permissible and critical

crack sizes

● Non-LEFM based analyses are ▬ essential for service failure FCG analyses

● Non-LEFM based and EBA analyses are▬ essential for test FCG analyses and their implications for se rvice aircraft

● Quantitative Fractography (QF) is essential in all cases

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DSTO examples of non-LEFM based FCG analyses

Aircraft / engine Component and materialLoad

history Service Test AnalysisVA

Boeing F/A-18 A/B

GD/LM F-111

P&W 125B

wing spar boom: AA 7075-T6

wing attachment boom: AISI 4340 steel

trailing edge flap lugs: AA 7050-T7451

vertical tail root: AA 7050-T7451

centre barrel bulkhead: AA 7050-T7451

rudder hinge fitting: AA 7075-T73

horizontal tail spindle: AF1410 steel

lower wing skin: 2024-T851

engine bearing cage: AISI 4340 steel

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

analytical

analytical

analytical

analytical

analytical

analytical

analytical

analytical

analytical

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● ●Aermacchi MB-326H


NLR examples of non-LEFM based FCG analyses

Aircraft / engine Component and materialLoad

history Service Test Analysis~ CA VA

Sikorsky S-61N

Northrop NF-5

Fokker F100

Fokker F100

BAC 1-11

Airbus A380

Mega Liner Barrel

(MLB)

NHIndustries NH-90

P&W F100-PW-220

rotor blade: AA 6061-T6

wing root lower skin: AA 7075-T651

vertical stabilizer rib: AA 7175-T736

fuselage lap splices: AA 2024-T3

fuselage lap splices: AA 2024-T3

window frame: AA 7175-T73

window panel: GLARE 3

passenger door beam: GLARE 3

tail hinge beam: AA 7075-T7351

air supply manifold support rod: In 718

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

hybrid

numerical

numerical

analytical

analytical

numerical

numerical

numerical

numerical

numerical

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*

*GLAss REinforced AA 2024-T3 laminates

*

● ●


FCG analysis examples


DSTO Non-LEFM example: Aermacchi MB -326H left wing loss and crash into the sea (1990)

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A7 - 076 during a training flight


Failure of left wing spar boom

● Lower spar boom failed adjacent to the end of the wing attachmentfitting during a high-g manoeuvre and at only 70% of the safe life


Inboard fracture surfaces of lower spar boom: aluminium alloy AA 7075-T6 extrusion

2 cm

Fatigue crack from abadly drilled fastenerhole:

Note fracture surfaceprogression markers.These were importantfor QF measurementsof crack growth


Left spar boom FCG tracked by correlating markers with G -meter data

5 mm

crac

kde

pth

(mm

)

flight number


Right wing spar boom FCG tracked by correlating progression markers with G -meter data

QF had to account for the complicationof tensile crack extensions (darker greyareas) due to peak manoeuvre loads


QF plots for the right wing spar boom crack: approximately exponential crack growth

crac

k de

pth

(mm

)

crac

k de

pth

(mm

)

flight number flight number

0Log(cd) (fn) Log(cd )= λ +


Intermediate status of the investigation

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� Physical causes of failure▬ badly drilled hole fatigue▬ high local stresses

� Contributing cause▬ poor quality control

� QF evidence for approximately exponential early fat igue crack growth:

a = a0eλN and ℓn(a) = λN + ℓn(a0)

where a is the crack size at time N, a0 is the initial crack size, and λ is a constant for a particular load history

N0a a eλ= ( ) ( )0Ln a N Ln a= λ +


Further actions (remedial measures)

� Non-Destructive Inspection (NDI) to detect cracks in fleet wings

� Teardown and QF for 9 crack-containing wings

� Reassessment of fleet wing lives

� Consider options for restoring/maintaining operational c apability▬ wing refurbishments

▬ wing replacements

� Final reassessment of fleet wing lives


Teardown results for spar cracks

service failure

Fraction of fatigue test life

Cra

ck D

epth

(m

m)

● 103 cracks found from examiningabout 1000 holes

● Cracks grew approximatelyexponentially: D17 hole cracks,including the service failure froma severe initial defect, grew faster

● Average trend for all cracks usedto reassess fleet wing lives: only 11 out of 69 aircraft fit for service!

I I I I I


Options for restoring/maintaining operational capab ility and final reassessment of fleet wing lives

� Wing refurbishment programme would require▬ NDI of thousands of fastener holes

▬ reaming and fitting oversize fasteners

● Wing replacement

▬ purchase sufficient wings to allow continued fleet operation: 30 new wing sets

● Final reassessment of fleet wing lives:

▬ initial quality of 2 crashed ‘new’ wings showed ove rall similarity to older wings

▬ thus the teardown results could be used to assess t he safe lives of new wings

▬ NDI to detect (any) cracks until fleet replacement by BAE Hawk Mk 127 aircraft

But estimated additional life unsatisfactory: programme rejected


NLR hybrid example : Sikorsky S -61N PH-NZCrotor blade loss and crash into the sea (1974)

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Aircraft type (PH-NZD)


blade 3fracture

PH-NZC recovery from the North Sea


Blade 3 fracture: 6061-T6 aluminium alloy(AlMgSi, peak aged) extruded spar

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phases of the spar fatigue life1 – initiation: corrosion pits + fatigue2 – thickness penetration3 – pocket-sealed flat fatigue fracture4 – flat fatigue fracture: pressure loss “detectable” 5 – slant fatigue and final fracture

25 mm


Blade 3 fracture: fatigue striation spacings for ph ase 4 along the spar (see previous slide); SEM = scanning elect ron microscopy, TEM = transmission electron microscopy

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● Trend line integration from a p to a fs gives N = 24,957 cycles

● Best fit QF trend indicates exponential FCG rate be haviour


Predicting S-61N detectable FCG lives (I): flow cha rts of calculations using best fit QF trend line and desig n data

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LEFM ‘Paris Law’● best fit striation spacings give:ΔKap = 24.3MPa√m; ΔKafs = 33.2MPa√m

● use striation spacings and these ΔK valuesto obtain reference ‘law’:

● Integration equation:

where ap and afs are effective crack lengths

AssumptionIAS either 120 or 130 knots

Calculate sets of ap and afs for: ● N = 24,957 cycles● ΔS at 120 and 130 knots

Calculate N for lower IAS using the two sets of ap and afs values

Convert N into flight hours● rotor speed 203 rpm● 1 FCG cycle per revolution

da / dN = 15 62.365 10 ( S a)−× ∆ π

−= × 12 6N 6.818 10 (ΔS) − 2 21/ a 1/p fsa


Predicting S-61N detectable FCG lives (II): spar ou tboard FCG lives from pocket edge, a p, to full slant fracture, a fs, as f (IAS)

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● IAS reduction to 90 knots gives FCG minimum of 12 hours = 4 maximum flight time.This IAS reduction became mandatory after in-flight detect ion of a pressure loss

×


Causes of failure and remedial actions

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● Physical cause: spar/pocket disbond corrosion p its fatigue

● Contributing causes▬ spar/pocket disbonds allowed within limits▬ highly inadequate design fatigue testing and analys is▬ high working stresses at 120–130 knots high crack gr owth rates▬ blade-mounted crack detection pressure gauges reada ble only when rotor stationary

� Remedial actions▬ detected spar/pocket disbonds immediately sealed; repair within 2 months▬ cockpit (in-flight) crack detection pressure gauges▬ in-flight pressure loss indication maximum spee d 90 knots to greatly

reduce working stresses and give detectable FCG lives more t han 4the maximum flight time

×


FCG analysis: some final remarks


Service and full-scale test FCG aspectsto be considered and /or accounted for

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● CA, R ● VA, R

▬ spectra and load histories▬ peak loads, underloads▬ short and long load sequences

● Geometry, stress concentrations● Secondary bending, load-shedding, residual stresses● Initial discontinuity sizes (effective crack sizes)

● Environment + cycle frequency effects (service)● Fretting effects on short crack growth● Differences in CA and VA fracture topography

Loads and stresses

Components and structures

Miscellaneous


FCG Quantitative Fractography (QF)

� Nominally constant amplitude (CA)

● Measurements of striation spacings versus a to obtain da/dN versus a plots▬ derive a versus N plots numerically, or analytically by equation fitting

● Variable amplitude (VA)

● Measurements of frequent or infrequent markers (inherent or imposed)versus a to obtain partial a versus N plots

▬ check for an analytical fit to “complete” the a versus N plots

● Measurements of frequent marker spacings (natural or imposed) – andmaybe striation spacings – versus a to obtain da/dN versus a plots

▬ derive a versus N plots numerically, or analytically by equation fitt ing:these can give insights into e.g. the effects of pe ak loads

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Nominally constant amplitude (CA) FCG

Variable amplitude (VA) FCGload history

load history

stress

stress

Note: Choice of VA measurement methods depends on t he marking characteristics Note: Choice of VA measurement methods depends on t he marking characteristics


Bibliography


Bibliography - I

� Barter, S.A. (2011), The quantification of fatigue crack in itiators in aluminium alloy 7050-T7451using quantitative fractography, International Journal of Structural Integrity, 2(3), 243-263.

� Barter, S. A., Lynch, S.P. and Wanhill, R.J.H. (2011), Failu re analysis of metallic materials: a shortcourse, NLR Technical Publication NLR-TP-2010-550, Natio nal Aerospace Laboratory NLR,Amsterdam, the Netherlands.

� Barter, S., McDonald, M. and Molent, L. (2005), Fleet fatigu e life interpretation from full-scale andcoupon fatigue tests - a simplified approach, USAF Structur al Integrity Program (ASIP) Conference2005, Memphis, USA.

� Barter, S.A., Molent, L. and Wanhill, R.J.H. (2012), Typica l fatigue-initiating discontinuities inmetallic aircraft structures, International Journal of Fatigue, 41(1), 11−22.

� Barter, S.A. and Wanhill, R.J.H. (2008), Marker loads for qu antitative fractography (QF) of fatigue inaerospace alloys, NLR Technical Report NLR-TR-2008-644, N ational Aerospace Laboratory NLR,Amsterdam, the Netherlands.

� Brooks, C. R. and Choudhury, A. (2001), Failure Analysis of Engineering Materials, McGraw-Hill,New York, USA.

� Buntin, W.D. (1977), Application of fracture mechanics to t he F-111 airplane, in: AGARDConference Proceedings No. 221 on Fracture Mechanics Design Methodology, pp. 3-1 – 3-12,Advisory Group for Aerospace Research and Development, Neu illy-sur-Seine, France.

� Campbell, G. S. and Lahey, R. (1984), A survey of serious airc raft accidents involving fatiguefracture, International Journal of Fatigue, 6(1), 25−30.

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Bibliography - II� Carpinteri, A. and Paggi, M. (2009), A unified interpretati on of the power laws in fatigue and the

analytical correlations between cyclic properties of engi neering materials, International Journal ofFatigue, 31, 1524-1531.

� Eijkhout, M.T. (1994), Fractographic analysis of longitud inal fuselage lapjoint at stringer 42 ofFokker 100 full scale test article TA15 after 126350 simulat ed flights, Fokker Report RT2160,Fokker Aircraft Ltd., Amsterdam, the Netherlands.

� Findley, S. J. and Harrison, N. D. (2002), Why aircraft fail, Materials Today, 5(11), 18−25.

� Hu, W., Tong, Y.C., Walker, K.F., Mongru, D., Amaratunga, R. and Jackson, P. (2006), A review andassessment of current lifing methodologies and tools in Air Vehicles Division, DSTO ResearchReport DSTO-RR-0321, DSTO Defence Science and Technology O rganisation, Fishermans Bend,Victoria 3207, Australia.

� Irving, P.E., Lin, J. and Bristow, J.W. (2003), Damage toler ance in helicopters, Report on the RoundRobin challenge, in: American Helicopter Society 59th Annual Forum, 6-8 May 2003, Phoenix, AZ,MIRA Digital Publishing, St. Louis, MP, USA.

� Iyyer, N., Sarkar, S., Merrill, R. and Phan, N. (2007), Aircr aft life management using crack initiationand crack growth models – P-3C aircraft experience, International Journal of Fatigue, 29, 1584-1607.

� Molent, L., Barter, S.A. and Wanhill, R.J.H. (2010), The lea d crack fatigue lifing framework, DSTOResearch Report DSTO-RR-0353, DSTO Defence Science and Tec hnology Organisation,Fishermans Bend, Victoria 3207, Australia.

� Molent, L., Barter, S.A. and Wanhill, R.J.H. (2011), The lea d crack fatigue lifing framework,International Journal of Fatigue, 33, 323-331.

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Bibliography - III

� Molent, L., McDonald, M., Barter, S. and Jones, R. (2008), Ev aluation of spectrum fatigue crackgrowth using variable amplitude data, International Journal of Fatigue, 30, 119-137.

� Redmond, G. (2001), From 'safe life' to fracture mechanics - F111 aircraft cold temperature prooftesting at RAAF Amberley, in: Proceedings of the 10th Asia-Pacific Conference on Non-DestructiveTesting, 17-21 September 2001, Brisbane, Australia.

� Tiffany, C.F., Gallagher, J.P. and Babish IV, C.A. (2010), T hreats to structural safety, including acompendium of selected structural accidents/incidents, U SAF Technical Report ASC-TR-2010-5002, Aeronautical Systems Center Engineering Directorat e, Wright-Patterson Air Force Base, OH45433-7101, USA.

� Wanhill, R.J.H. (2003), Material-based failure analysis o f a helicopter rotor hub, Practical FailureAnalysis, 3(2), 59-69.

� Wanhill, R. J. H. (2009), Some notable aircraft service fail ures investigated by the NationalAerospace Laboratory (NLR), Structural Integrity and Life, 9(2), 71−87.

� Wanhill, R.J.H. and Hattenberg, T. (2006), Fractography-b ased estimation of fatigue crack“initiation” and growth lives in aircraft components, NLR Te chnical Publication NLR-TP-2006-184,National Aerospace Laboratory NLR, Amsterdam, the Netherl ands.

� Wanhill, R.J.H., Hattenberg, T. and Van der Hoeven, W. (2001 ), A practical investigation of aircraftpressure cabin MSD fatigue and corrosion, NLR Contract Repo rt NLR-CR-2001-256, NationalAerospace Laboratory NLR, Amsterdam, the Netherlands.

� Wanhill, R.J.H., Molent, L. and Barter, S.A. (2012/2013), F racture mechanics in aircraft failureanalysis: uses and limitations, Engineering Failure Analysis, available online 15 November 2012.

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G’DAY MATE. THEY’RE SAFE – THEY’VE BEENLEFM ANALYSED AND

TESTED

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