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2007 Aircraft Structural IntegrityProgram ConferenceProgram Conference
The Effect of StressThe Effect of Stress Intensity Factor Models on Inspection Intervalson Inspection Intervals
Lt Col Scott FawazCenter for Aircraft Structural Life Extension
United States Air Force Academy
I n t e g r i t y - S e r v i c e - E x c e l l e n c e
y
DISTRIBUTION STATEMENT A: Approved for public release: distribution is unlimited
Acknowledgements
• Dr. Börje Andersson - Swedish Defense Research AgencyResearch Agency
• Daniel Hill – CAStLE Research Engineer• Dr. R. A. Saravanan – CAStLE Metallurgist• Jim Harter – AFGROW Lead Engineerg• Alex Litvinov – AFGROW Software Engineer
Outline• K Solutions
° Geometric & Loading Parameter SpaceGeometric & Loading Parameter Space° Verification° Validation
F ti Lif P di ti U i N K• Fatigue Life Predictions Using New KSolutions° Fatigue Life° Fatigue Life° Continuing Damage Scenario
› Phase I LifeC k Si› Crack Size
° Effect of r/t
• ConclusionsConclusions
Small differences in K SolutionsSmall differences in K Solutionsyield large cumulative differences
i f i lifin fatigue life
…and large differences in K solutions yield even a larger cumulative difference in fatigue life
Parameter SpaceK-Solutions, ≈ 1.0 million CPU Hours
• Geometry° Centrally Located Straight Shank Hole° 0.1 ≤ r/t ≤ 10.0
σbendingσbypass
› 0.1, 0.111, 0.125, 0.1428, 0.1667, 0.2, 0.25, 0.333, 0.5, 0.667, 0.75, 0.8, 1.0, 1.25, 1.333, 1.5, 1.667 2.0, 3.0, 4.0, 5.0, 6.0, 7.0, 8.0, 9.0, 10.0 (r/t = 0.5, 1.0)
° Finite Width/Height Plate/h 0 0025
Pcos2θ
› r/h = 0.0025› r/b = 0.0025
• Crack Shapes° 0.1 ≤ a/c ≤ 10.0
› 0 1 0 111 0 125 0 1428 0 1667 0 2 c1
2h
2rc2› 0.1, 0.111, 0.125, 0.1428, 0.1667, 0.2, 0.25, 0.333, 0.5, 0.667, 0.75, 0.8, 1.0, 1.25, 1.333, 1.5, 2.0, 3.0, 4.0, 5.0, 6.0, 7.0, 8.0, 9.0, 10.0 (a/c = 0.2, 0.5, 0.8, 1.0, 2.0)
° 0.1 ≤ a/t ≤ 0.99› 0.1, 0.2, 0.3, 0.4, 0.5, 0.6, 0.7, 0.8, 0.9,
2b
12
σ, , , , , , , , ,0.95, 0.99 (a/t = 0.2, 0.5, 0.8)
• Load Conditions° Tension° Bending
σo
a
σbending
g° Pin Loading (Bearing)
• 5,672,700 solutionsa2
a1
K S l tiK-SolutionVerificationVerification
Convergence: Shallow Crack1.0
/t 1 0
σo
0.8
r/t = 1.0a/c = 0.1a/t = 0.12b = 2h = 200r = 400
2b
2h
0.6
c2r
σoa
0 2
0.4
p = 2
p = 3
0.0
0.2 p = 4
p = 5
p = 60.0
0.0 0.2 0.4 0.6 0.8 1.0
Convergence: Deep Crack1.2
p = 2
3r/t = 1.0a/c = 0 1
0 8
1.0p = 3
p = 4
p = 5
p = 6
a/c = 0.1a/t = 0.992b = 2h = 200r = 400
0.6
0.8 p
σo
0.4
2b
2h
0 0
0.2
c2r
σo
a
0.00.0 0.2 0.4 0.6 0.8 1.0
K S l tiK-SolutionValidationValidation
Test Specimen Configurationσ
WW
Crack plane
ta
σa1
a2
c1 c2
Marker Load Spectrum
100 cycles 100 cycles 100 cycles
10 cycles 10 cycles 10 cycles
stre
ss 2000Cycles
2000Cycles
2000Cycles repeat2000
Cycles
1 program = 8170 cycles
Fatigue Life Prediction
3.5Test 3, 7075-T651w=49.17 mm, t=6.36 mm, r=3.23 mmc1=2.76 mm, c2=1.75 mm
3
)
c1 2.76 mm, c2 1.75 mm
2.5
k Le
ngth
(mm
)
AFGROW Lef t Crack
Measured Lef t Crack
2
Cra
ck Measured Lef t Crack
AFGROW Right Crack
Measured Right Crack
1.562,000 64,500 67,000 69,500 72,000 74,500
Cycles
Crack Shape Development
Crack Shape Development
Crack Shape Development
F ti Lif P di ti U iFatigue Life Predictions Using New K Solutions
Geometry for Assessing Effect on LifeSmall Crack – Thin Sheet
W = 1.14 in, t = 0.063 in, D = 3/16 in
a 0 01 in c 0 01 in a /t 0 2
σo σbending
ai = 0.01 in, ci = 0.01 in, ai /t = 0.2
ai /ci = 1.0, r/t = 1.5
TSR = 1.0, BSR = 0.4,Small Crack – Thick Sheet
W = 4.53 in, t = 0.25 in, D = ¾ in
0 05 i 0 05 i /t 0 2 2ai = 0.05 in, ci = 0.05 in, ai /t = 0.2
ai /ci = 1.0, r/t = 1.5
TSR = 1 0 BSR = 0 4W
c12rc2
Large Crack – Thin Sheet
W = 1.14 in, t = 0.063 in, D = 3/16 in
0 05 i 0 05 i /t 0 8
TSR 1.0, BSR 0.4
σoσbending
ai = 0.05 in, ci = 0.05 in, ai /t = 0.8
ai /ci = 1.0, r/t = 1.5
TSR = 1.0, BSR = 0.4
a2 a1t
Effect on Life – Small Crack, Thin Sheet60 Two Symmetric Corner Cracks
Single Corner CrackW = 1.14 in, t = 0.063 in, D = 3/16 inai = 0.01 in, ci = 0.01 inai /t = 0.2, ai /ci = 1.0, r/t = 1.5TSR = 1 0 BSR = 0 4
40
50
(%)
TSR = 1.0, BSR = 0.4
30
20
10
0CA 10 CA 18 Falstaff 10 Falstaff 18 TWIST 10 TWIST 18 Marker 10 Marker 18
Spectrum Type and Maximum Stress Level
Effect on Life – Small Crack, Thick Sheet
45
50Two Symmetric Corner CracksSingle Corner Crack
W = 4.53 in, t = 0.25 in, D = 3/4 inai = 0.05 in, ci = 0.05 inai /t = 0.2, ai /ci = 1.0, r/t = 1.5TSR = 1.0, BSR = 0.4
35
40
(%)
TSR 1.0, BSR 0.4
20
25
30
10
15
20
0
5
CA 10 CA 18 Falstaff 10 Falstaff 18 TWIST 10 Marker 10Potential for initial inspection of damage tolerant (rogue flaw)
i l h i h i f lif-5
CA 10 CA 18 Falstaff 10 Falstaff 18 TWIST 10 Marker 10
Spectrum Type and Maximum Stress Levelnot occurring early enough in the aircraft life
Effect on Life – Large Crack, Thin Sheet300
Two Symmetric Corner CracksSingle Corner Crack
W = 1.14 in, t = 0.063 in, D = 3/16 in
200
250
(%)
ai = 0.05 in, ci = 0.05 inai /t = 0.8, ai /ci = 1.0, r/t = 1.5TSR = 1.0, BSR = 0.4
150
100
50
Potential for initial inspection of damage tolerant (rogue flaw) i l h i h i f lif
0CA 10 CA 18 Falstaff 10 Falstaff 18 TWIST 10 TWIST 18 Marker 10 Marker 18
Spectrum Type and Maximum Stress Level
not occurring early enough in the aircraft life
Geometry for Assessing Effect on Continuing Damage Scenario σg σo σbending
W = 1.14 in, t = 0.063 in, D = 3/16 in
0 05 i 0 05 ia1 = 0.05 in, c1 = 0.05 in
a2 = 0.005 in, c2 = 0.005 in
a1 /t = 0.8, a2 /t = 0.08, ai /ci = 1.0, r/t = 1.5 2a1 /t 0.8, a2 /t 0.08, ai /ci 1.0, r/t 1.5
TSR = 1.0, BSR = 0.4W
c12rc2
σo
σbending
a2 a1 t
Effect on Continuing Damage ScenarioPhase I Life
350
400W = 1.14 in, t = 0.063 in, D = 3/16 ina1 = 0.05 in, c1 = 0.05 ina2 = 0.005 in, c2 = 0.005 ina /t = 0 8 a /t = 0 08 a /c = 1 0 r/t = 1 5
300
(%)
a1 /t = 0.8, a2 /t = 0.08, ai /ci = 1.0, r/t = 1.5TSR = 1.0, BSR = 0.4
200
250
100
150
50
100
0CA 10 CA 18 Falstaff 10 Falstaff 18 TWIST 10 TWIST 18 Marker 10 Marker 18
Spectrum Type and Maximum Stress Level
Effect on Continuing Damage ScenarioPhase I Crack Lengthg
60a-Crack Tip
c-Crack TipW = 1.14 in, t = 0.063 in, D = 3/16 ina1 = 0.05 in, c1 = 0.05 ina2 = 0.005 in, c2 = 0.005 ina /t = 0 8 a /t = 0 08 a /c = 1 0 r/t = 1 5
40
50
(%)
a1 /t = 0.8, a2 /t = 0.08, ai /ci = 1.0, r/t = 1.5TSR = 1.0, BSR = 0.4
30
20
10
0CA 10 CA 18 Falstaff 18 TWIST 10 TWIST 18 Marker 10 Marker 18
Spectrum Type and Maximum Stress Level
Effect of r/t – Symmetric Corner Cracks
35
40 Symmetric Corner CracksW/D = 100, ai/ci = 1.0ai = ci = 0.05 mm"Thin Sheet" → a/t = 0.8
30F/A Thin SheetN/R Thin SheetF/A Thick Sheet
"Thick Sheet" → a/t = 0.2
20
25
Nr/t
= 3
.0
(%) N/R Thick Sheet
10
15Ni/
N
5
10
00 0.5 1 1.5 2 2.5 3
r/t
Effect of r/t – Single Corner Crack
90
100 Single Corner CrackW/D = 100, ai/ci = 1.0ai = ci = 0.05 in"Thin Sheet" → a/t = 0.8
70
80F/A Thin SheetN/R Thin SheetF/A Thick Sheet
"Thick Sheet" → a/t = 0.2
50
60
Nr/t
= 3
.0
(%) N/R Thick Sheet
30
40Ni/
N
10
20
00 0.5 1 1.5 2 2.5 3
r/t
Conclusions
• Verification° hp-version FEA + Splitting Scheme = Accurate K-
Solutions• Validation• Validation
° Fatigue life predictions are slightly conservative• 5 672 700 K solutions for unsymmetric corner5,672,700 K solutions for unsymmetric corner
cracks at a hole subject to tension, bending, bearing° Solutions available in tabular form – currently in
AFGROW› 75 – 1.5MB ASCII files
° Source code for multi-dimensional interpolation also available
Significance• Single vs. Double Cracks
° Difference always larger for single cracksy g g
• Effect on Fatigue Life° Small cracks in thin sheets: 20-50%
S ll k i thi k h t 25 45%° Small cracks in thick sheets: 25-45%° Large cracks in thin sheets: 90-300%° Continuing damage scenario: 125-350%
• Effect on Inspections° Possibility of initial inspection not early enough in aircraft life° Possibility of recurring inspections not occurring as frequently as° Possibility of recurring inspections not occurring as frequently as
required
• Effect of r/t° Significant for large cracks in thin sheets° Negligible for small cracks in thick sheets
