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