Small Crack BehaviorSmall-Crack Behavior

1ffa

Motivation

Observations on the fatigue behavior of engineered g gmaterials support the concept that “crack propagation” from micro-structural discontinuities consume a largeportion of the fatigue lives from LCF to HCF:portion of the fatigue lives from LCF to HCF:

- Aluminum alloys (2024-T3, 7075-T6 and T651)- Titanium alloy (Ti-6Al-4V engine disks)- Magnesium alloy (AZ91E)- Steel (4340)

2

Design Concept Using Small-Crack Theory

Flaw size

DSmall-

None 1 m 10 m 0.1 mm 1 mm 1 m

Safe-life Durability Damagetolerance

Fail-safecracktheory

o Economic- life extension

o Inspectable flaw size

o Redundant structure

o Hidden or uninspectable structureo Manufacturing- and

i i d d d service-induced damageo EIFS to fit S-N or -N behavior

3

OUTLINE OF PRESENTATION

• A fracture-mechanics approach to fatigue-life prediction

• Application of small-crack theory to panels exposedto laboratory air and to an aggressive environmenty gg

• Application of small-crack theory to coupons exposedand pre-corroded at several U.S. Air Force basesand pre corroded at several U.S. Air Force bases

• Application of small-crack theory to an engine material

• Concluding remarks

4

AGARD / NASA / CAE SMALL-CRACK PROGRAMS1984-1994

AGARD R-732, R-767 – P.R. Edwards and J.C. Newman, Jr.AGARD R-766 – A.J. Mom and M.D. RaizenneAGARD R-766 (Addendum) – T. Pardessus, E. Jany and M.D. RaizenneNASA/CAE RP-1309 – J.C. Newman, Jr. and X.R. Wu

Materials Loading Actions*2024-T3 Constant amplitude (R = -2 to 0.5)7075 T6 FALSTAFF (fighter spectra)7075-T6 FALSTAFF (fighter spectra)Lc9CS (7075-T6 clad) Inverted FALSTAFF2090-T8E41 Gaussian (R = -1 random spectra)4340 Steel TWIST (transport spectra4340 Steel TWIST (transport spectraTi-6Al-4V Mini-TWISTIMI-685 Felix-28 (helicopter spectra)Ti-17 Turbistan (engine spectra)( g p )

* Loading and materials are not associated5

SMALL - CRACK TEST SPECIMEN

6

EFFECT OF POLISHING ON NOTCH SURFACES

7

STRESS LIFE CURVES UNDER SPECTRUM LOADINGNewman and Edwards 1988

350 FALSTAFF 2024-T3

Newman and Edwards, 1988

250

300B = 2.3 mm SENTKT = 3.17Failure

Smax MPa 150

200

250

100

1502a = 20 mm

0.2 mm 2.3 mm

m

0

50

103 104 105 106 107

N, cycles10 10 10 10 10

8

CRACK INITIATION SITESIN 7075 T6 AND LC9cs ALLOYSIN 7075-T6 AND LC9cs ALLOYS

(a) 7075-T6 (b) LC9cs (7075 clad)(a) 7075-T6 (b) LC9cs (7075 clad)

9

CRACK GROWTH IN CLAD AND BARE ALLOYS

10

CORNER-CRACK GROWTH IN LC9cs CLAD ALLOY

Clad layer Crack

First crack atN = 5000 N = 8000 N = 14000 cycles

(Specimen failure in 28000 cycles)( p y )

11

TYPICAL SMALL- AND LARGE-CRACK GROWTH RATE DATA

Constant-amplitude loadingR = constant

S1 < S2 < S3

Small crack Large crack

S3

1 2 3

Small crack

da , dcdN dN__ __

S3

Large crackS1

S2

(K decreasing test)

Steady state

S1

Kth K

12

SMALL- AND LARGE-CRACK GROWTH RATES IN 7075-T6

1e-3 Keff7075-T6 [23]KT = 3.15R = -1

da/dN

1e-4R 1

FASTRAN ( = 1.8)ai = ci = 6 m

or dc/dNmm/cycle

1e-6

1e-5PhillipsLarge cracks(K; dc/dN)

i i

1e-7

1e-6 (K; dc/dN)

Small surfacecracks at notch (K)

1e-8

Smax = 80 MPa

0 5 1 2 5 10 20 50K or Keff, MPam

0.5 1 2 5 10 20 50

13

SMALL- AND LARGE-CRACK GROWTH RATES IN 2024-T3

1e-3Keff

2024-T3 [17]KT = 3.15R = 0

da/dNor 1e 5

1e-4

Small surfacecracks at notch (K) or

dc/dNmm/cycle

1e-6

1e-5

FASTRAN ( = 2)

Smax = 110 MPa

1e-7

1e 6

Phillips [33]Large cracks

ai = ci = 6 m

1e-80 5 1 2 5 10 20

g(K; dc/dN)

K or Keff, MPa-m1/20.5 1 2 5 10 20

14

MEASURED AND PREDICTED FATIGUE LIVES FOR 2024-T3SPECIMENS UNDER UNIFORM STRESS

500 u 2024-T3 [34]B = 2.3 mmw = 25 4 mm

Hardrath et al.

400

ys

w = 25.4 mmKT = 1

R = 0

Smax, MPa

200

300R = -1

100

200

FASTRANai = ci = 20 m

102 103 104 105 106 107 1080

100 ai = ci = 20 m

Nf, cycles

102 103 104 105 106 107 108

15

MEASURED AND CALCULATED FATIGUE LIVES FOR 2024-T3SPECIMENS WITH CIRCULAR HOLE

400 Hardrath et al.2024-T3

300

B = 2.3 mmr = 1.6 mmw = 25.4 mmKT = 3

Smax MPa 200

KT 3

R = 0R = -1

100

Closure modelai = ci = 6 m

Elastic

1e+2 1e+3 1e+4 1e+5 1e+6 1e+70

Elastic-plastic

Nf, cycles

e e 3 e e 5 e 6 e

16

SOME FLIGHTS IN TWIST TRANSPORT WING SPECTRUM

100

60

80

Appliedstresses 40

60

20

0 200 400 600 800 1000 1200 1400-20

0

Time

0 200 400 600 800 1000 1200 1400

17

MEASURED AND PREDICTED FATIGUE-LIFE BEHAVIOR UNDER MINI-TWIST SPECTRUMO U S S C U

300 Test [22,27]Mini-TWIST Test (KT = 3.15)

2507075-T6 BareLC9cs Clad

Smax,MPa 150

200

MPa

100LC9cs or 7075-T6B = 2 or 2 3 mm

FASTRAN6

0

50B = 2 or 2.3 mmw = 50 mmr = 3.18 mmKT = 3.15

ai = ci = 77 mai = ci = 6 m

Nf, cycles105 106 107

0

18

Major Result of Small-Crack Programsj g

Fatigue is “crack propagation” from micro-structural

features or discontinuities (such as inclusion-particles,

voids, and cladding) for many engineering materials.

19

OUTLINE OF PRESENTATION

• A fracture-mechanics approach to fatigue-life prediction

• Application of small-crack theory to panels exposedto laboratory air and to an aggressive environmenty gg

• Application of small-crack theory to coupons exposedand pre-corroded at several U.S. Air Force basesand pre corroded at several U.S. Air Force bases

• Application of small-crack theory to an engine material

• Concluding remarks

20

Influence of Manufacturing Procedures on Fatigue Lives of Riveted Single Lap Joints, Hartman, NLR, 1968

250 mm

1020

DD Rivet2r = 3.2 mm

S S160 mm

10 20 10S S

Rivet head (D)B = 1 mm

21

Cracks at Riveted Fastener Hole under Various Loading

S

M

Sb2 r

P

c

B

c

(a) Through crack

M

ca

B2 r

M

(b) Corner crackS S

c

Sp Sb+

22

Measured and Calculated Fatigue Lives on Riveted Lap Joints under Constant-Amplitude Loading

70 2024-T3 AlcladB = 1 mmHartman

50

60 Sm = 68.6 MPa

FASTRANai = ci = 6 m

SaMPa

30

40i i

= 0 = 5.8 m

20D = 5.0 to 5.2 mm

10 Test Series

1e+4 1e+5 1e+6 1e+7 1e+80

10 10 Test Series

Number of Load Cycles

1e+4 1e+5 1e+6 1e+7 1e+8

23

Panel Test Program Conducted by Furuta, Terada, and Sashikuma, KHI and NAL, Japan, ICAF ‘97

Test conditions:• 100O countersunk rivets

• Ambient (laboratory air and room temperature)

• Salt-water solution (3.5% NaCl immersed)

24

TEST METHODP I

25

Small- and Large-Crack Growth Rates on 2024-T3 Bare Aluminum Alloy Sheet

1e-2 2024-T3 (Piascik and Willard)Deaerated 1% NaCl (-700 mVSCE)R (f = 5 Hz)

1e-4

1e-3 R (f = 5 Hz)0.050.700.750.77

da/dNmm/cycle 1e-5

0.770.80

1e-6 2024-T3 Bare & AlcladLab air (f = 10 Hz)

1e-8

1e-7

Keff, MPam

1e 80.5 1 2 5 10 20

26

Measured and Predicted Fatigue Crack Growth in LapJoint Panels under Ambient and Salt-Water Conditions

1000Furuta et al (1997)2024-T3Smax = 96 MPa; R = 0.125

100 AmbientCorrosive (3.5% NaCl)

Cracklength, c, mm

1

10

0.1

1 FASTRAN

ai = ci = 6 m(No bending andno interference)

0.01

no interference)

N, cycles0 1e+5 2e+5 3e+5

27

Measured and Predicted Fatigue Lives for Type 1Lap-Joint Panels

1e+6 Type 1 tests (Furuta et al)FASTRAN (ai = ci = 6 m)

Nf 1e+5

Bending Bending No bending

Edge clamps: No No Yes Yes

1e+4

Environment: Ambient Corrosive Ambient Corrosive

28

Measured and Predicted Fatigue Lives for VariousTypes of Lap-Joint Panels

1e+6

Nf 1e+5

1e+4

Tests (Furuta et al)FASTRAN (ai = ci = 6 m)

Type: 3 4 3 4 3 4 2 3Environment: Ambient Corrosive Corrosive Corrosive

Tear strap: Yes Yes Yes No YesThickness (mm): 1 1 2 1Thickness (mm): 1 1 2 1

29

Tear-Down Examinations of Actual Aircraft Panels with Riveted Lap Joints

Fuselage Panels contain:Counterbore rivetsC t k i t

Panels 3 & 6Panels 3 & 6

Countersunk rivetsStraight-shank rivets

Panel 1

30

Comparison of Crack Growth Rate Data from Aircraft Panels and AGARD Small Crack Database

10-3

Small Crack Data(AGARD R 732)

10-5

10-4 (AGARD R-732)

F ti k

10-6

10Fatigue crackgrowth rate(mm/cycle)

Panel Data

10-7 Skin Thickness6 50 100 m

1 10 100 1000 1000010-8

Crack length (m)Crack length (m)

31

Prediction of Fatigue Crack Growth at Rivet Holes in Actual Aircraft Fuselage Joints

10050 m100 m 6 m

FASTRAN

10

Panel data

0 1

1Fatigue cracklength (mm)

0.01

0.1

0 25000 50000 75000 4000000.001

Fuselage pressure cyclesFuselage pressure cycles

32

OUTLINE OF PRESENTATION

• A fracture-mechanics approach to fatigue-life prediction

• Application of small-crack theory to panels exposedto laboratory air and to an aggressive environmenty gg

• Application of small-crack theory to coupons exposedand pre-corroded at several U.S. Air Force basesand pre corroded at several U.S. Air Force bases

• Application of small-crack theory to an engine material

• Concluding remarks

33

Corrosion Severity Distribution Across Air Force Bases - Aluminum Corrosion Rates

• Severity is 200:1 betweenBattelle (Abbott)

Severity is 200:1 between air bases

• Data are for boldly exposed coupons

• Relevance to aircraft:• Probably similar rates on

upper wing skin after CPC protection lostCPC protection lost

• Different/higher rates in crevices or lap joints

• Different/lower rates e e t/ o e atesinside aircraft

34

Three-Hole Bend Specimen Exposed and Tested under Reverse Bending (R = -1) Loads in the Laboratory

D = 0.25 in.Abbott (Battelle)

Area of applied load

B 0 16 i

W = 1.0 in.

B = 0.16 in.

35

Cumulative Distribution of Pit Depths for Various Air Bases and Months of Exposureand Months of Exposure

100 DAB (3 month)

80

DAB (6 month)DAB (9 month) DAB (12 month) WPAFB (3 month) WPAFB (6 month)

CumulativeDistribution

Function 40

60WPAFB (6 month)

Function

20

40

1 10 1000

d, m

36

Crack-Closure Model Correlates Rates for 2024-T3

10-4

10-3Hudson, Phillips & Dubensky2024-T3Middle crack tensionB = 2 3 mm

10-6

10-5B = 2.3 mm

= 1

10-8

10-7 dc/dNm/cycle

= 2R

0.70 5

10-10

10-9

0.50.30-1-2

1 10 10010-12

10-11

-2

1 10 100

Keff, MPa-m1/2

37

Measured and Calculated Fatigue Lives for Laboratory Coupons with No Exposure

40 Abbott (Battelle)2024-T3

Remote

30

B = 0.16 in.W = 1.0 inD = 0.25 in.Cantilever bending:R 1Remote

bendingstress,Sb, ksi

20 Lab air (no exposure)FASTRAN:

R = -1

10FASTRAN:ai = ci = 4e-5 in. (1 m)ai = ci = 8e-5 in. (2 m)ai = ci = 1 2e-4 in (3 m)

Nf cycles

103 104 105 106 107 1080

ai = ci = 1.2e-4 in. (3 m)

Nf, cycles

38

Measured and Calculated Fatigue Live for Coupons Exposed at Wright-Patterson Air Force Base

30WPAFB (3-month)WPAFB (6 month)

Remote 20

25WPAFB (6-month)EIFS = 4 mEIFS = 6 m

Remotebendingstress,Sb, ksi

15

5

10

Nf cycles1e+4 1e+5 1e+6 1e+7 1e+8

0

Nf, cycles

39

Measured and Calculated Fatigue Live for Coupons Exposed at Daytona Air Force Base

30 DAB (3-month)DAB (6 th)

R t 20

25DAB (6-month)DAB (9-month)DAB (12-month)EIFS = 12 m

Remotebendingstress, Sb, ksi

15

20EIFS = 40 mEIFS = 500 m

b

5

10

N l1e+4 1e+5 1e+6 1e+7 1e+8

0

Nf, cycles

40

Relationship between EIFS and Exposure Pit Depths at 50 Percentile for 2024-T3 Aluminum Alloy

1000Wright-PattersonDaytona

100

DaytonaPerfect agreement

EIFS = d5 + 0.005 d53

EIFS, m

10

d5

10

Pit Depth (50%) d m

1 10 1001

Pit Depth (50%), d5, m

41

OUTLINE OF PRESENTATION

• A fracture-mechanics approach to fatigue-life prediction

• Application of small-crack theory to panels exposedto laboratory air and to an aggressive environmenty gg

• Application of small-crack theory to coupons exposedand pre-corroded at several U.S. Air Force basesand pre corroded at several U.S. Air Force bases

• Application of small-crack theory to an engine material

• Concluding remarks

42

FATIGUE-LIFE PREDICTION METHOD BASEDON SMALL-CRACK THEORY IN AN ENGINE

MATERIAL

James C. Newman, Jr.Aerospace Engineering

Balkrishna S. AnnigeriStructures and DynamicsAerospace Engineering

Mississippi State UniversityMississippi State, MS

USA

Structures and DynamicsPratt & Whitney

East Hartford, CTUSAUSA USA

Research sponsored by: Pratt & Whitney 43

Material

Titanium alloy (Ti-6Al-4V stoa) forged blade material used in U.S. Air Force high-cycle-fatigue (HCF) test and analysisprogram: AFRL-ML-WP-TR-2001-4159 (Average size of primary alpha grain: 15 m)primary alpha grain: 15 m)

Metallic Materials Properties Development Standards(MMPDS) H db k J 2003(MMPDS) Handbook – January 2003

Newman, J.C., Jr., Ruschau, J.J. and Hill, M.R., “Improved , , , , , , pTest Method for Very Low Fatigue-Crack-Growth-Rate Data”, Fatigue Fract Engng Mater Struct Journal, 2011

44

Fatigue and Crack-Growth Specimens

Compact Specimen

Fatigue Specimens45

ASTM E-647 Load-Reduction Method Causes a Width Effect Ruschau and Newman (2008)Ruschau and Newman (2008)

Calculated

46

Comparison with Test Data from U.S. Air Force Report

Calculated

47

Compression Precracking Produces Width Independent DataRuschau and Newman (2008)usc au a d e a ( 008)

Calculated

48

Effective Stress-Intensity Factor Range against Rate Data

49

Crack-Growth Rates for Small- and Large-Cracks at R = 0.1

50

Crack-Growth Rates for Small- and Large-Cracks at R = 0.1

??

51

Crack-Growth Rates for Small- and Large-Cracks at Various R

52

Measured and Predicted Small-Crack Growth

53

Measured and Predicted Small-Crack Growth

??

Tests (average):

Nf = 1.8 x 106 cycles

Analysis:

Nf = 162,000 cycles

54

Measured and Predicted Small- and Large-Crack Growth Rates

??

55

Fatigue Tests on Round Bars at R = 0.1T t D t f AFRL TR 4159Test Data from AFRL TR-4159

D

56

Fatigue Tests on Rectangular Sheets at R = 0.1T t D t f MMPDS (2003)Test Data from MMPDS (2003)

57

Fatigue Tests on Flat Sheets with Hole at Various RT t D t f MMPDS (2003)Test Data from MMPDS (2003)

58

Fatigue Tests on Double-Edge Notch at R = 0.1 & -1

Test Data from MMPDS (2003)

59

Fatigue Tests on Double-Edge Notch at R = 0.5 & 0.8

Test Data from MMPDS (2003)

60

CONCLUDING REMARKS

• Fatigue in aluminum alloys, and other engineering materials, is essentially “crack propagation” from micro structural featuresmicro-structural features.

• FASTRAN is able to predict the growth of small d l k d t t lit d dand large cracks under constant-amplitude and

spectrum loading reasonably well, exceptions are materials that produce rough fatigue surfaces andaccumulate fretting debris.

• FASTRAN is able to predict (or calculate) the fatiguep ( ) glives of notched and un-notched specimens under constant-amplitude and spectrum loading, exceptionsare severe loads on materials that produce roughare severe loads on materials that produce rough fatigue surfaces and accumulate fretting debris.

61