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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. Brje 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

    bendingbypass

    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

    Pcos2

    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

    , , , , , , , , ,0.95, 0.99 (a/t = 0.2, 0.5, 0.8) Load Conditions

    Tension Bending

    o

    a

    bending

    g Pin Loading (Bearing)

    5,672,700 solutionsa2

    a1

  • K S l tiK-SolutionVerificationVerification

  • Convergence: Shallow Crack1.0

    /t 1 0

    o

    0.8

    r/t = 1.0a/c = 0.1a/t = 0.12b = 2h = 200r = 400

    2b

    2h

    0.6

    c2r

    oa

    0 2

    0.4

    p = 2

    p = 3

    0.0

    0.2 p = 4p = 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

    o

    0.4

    2b

    2h

    0 0

    0.2

    c2r

    oa

    0.00.0 0.2 0.4 0.6 0.8 1.0

  • K S l tiK-SolutionValidationValidation

  • Test Specimen Configuration

    WW

    Crack plane

    ta

    a1

    a2

    c1 c2

  • Marker Load Spectrum

    100 cycles 100 cycles 100 cycles

    10 cycles 10 cycles 10 cycles

    stre

    ss 2000Cycles

    2000Cycles

    2000Cycles repeat

    2000Cycles

    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 CrackAFGROW 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

    o bending

    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 = in

    0 05 i 0 05 i /t 0 2 2ai = 0.05 in, ci = 0.05 in, ai /t = 0.2ai /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

    obending

    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 a1 t

  • 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 g o bending

    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 1.5TSR = 1.0, BSR = 0.4

    W

    c12rc2

    obending

    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