Damage Tolerance Risk Assessment of T-38 Wing Skin Cracks...• T-38 NDI team of experts assembled...

Damage Tolerance Risk Assessment of T-38 Wing Skin Cracks

Presented at the 2005 USAF ASIP Conference

November 29, 2005

Joseph W. Cardinaland

Dr. Hal BurnsideSouthwest Research Institute®

San Antonio, TX

Dr. Michael P. BlinnT-38 ASIP Mgr.

Hill AFB, UT

2

Acknowledgements

• 1Lt Jason Lee, Hill AFB• Cliff Massey, Randolph AFB• Jim Lankford, SwRI (Metallurgy)• Dave Wieland, SwRI (DTA & FEA)• Kurt Schrader, SwRI (Stress Spectra)

Work performed underSwRI Project No. 18-08760

USAF Contract No. F4260-00-D-0037, D.O. 0048for

Ogden Air Logistics CenterHill Air Force Base, UT

3

Presentation Outline

• Background• Objectives and Approach• Metallurgical Failure Analysis• Original and Repaired Geometry• FEA and DTA Results• Motivation for Risk Analyses• Development of EIFS Distribution• Risk Analyses and Sensitivity Studies• Conclusions and Recommendations• USAF Actions and Way Ahead• References

4

Background

• In March 2004, unusually large cracks detected and reported to the T-38 SPO/Engineering

• Affected aircraft flying a relatively severe mission

• Cracks found on the lower wing skin (LWS), a fatigue sensitive structural part of the T-38’s -29 wing

• The subject cracks were in a frequently inspected area of the LWS (i.e., 125 hour recurring intervals)

• Concern as to why these cracks grew so large in this critical structural area

5

Background-29 Wing Structure and FCLs

Note: Wing Fatigue Critical Locations (FCLs) Shown (e.g., A-1, A-4, etc.)

A-24

A-23 A-22A-18

A-20A-9

A-10a

A-21

A-12b A-12a

A-4

A-5

A-19

A-15A-17

A-1

A-8a

Area of Large Cracks

6

Background

RHS Wheel Well for –29 Wing

Area of Large Cracks

7

Background

Relatively large cracks (1.5 and 0.75 inch) were found on the LWS of T-38 wings near the 1.5 inch radius (WS 64.8 / 44% spar) at a known fatigue critical location (FCL A-9)

Dye Penetrant Inspection of FCL A-9 Wing Crack (TN 68-8201)

USAF Wing Inspection(Surface Eddy Current)

Crack

8

BackgroundFatigue Cracks at FCL A-9

TN 68-8201

TN 64-3261

Data:• Land Thickness: 0.147 inch• Wing Skin Thickness: 0.335 inch• Material: 7075-T7351 plate

• Each crack originated at inside(upper) corners of the LWS radius on the land (white arrow)• No defect or “rogue” flaw wasable to be identified

9

BackgroundActions Taken

• Detailed LWS inspections were immediately performed by the USAF at selected bases

• T-38 NDI team of experts assembled to assist in the inspection process

• Wing cracking summary:– 22 wings found with cracks– 11 wings condemned– 11 remaining wings were thought repairable

• Relatively large number of condemnations at one time affected -29 wings in inventory (i.e., this was an unanticipated event)

• Need to investigate risk before and after repair

10

Objectives & Approach

• Engineering topics considered:– Metallurgical failure analysis– Finite Element Analysis (FEA) of original & repaired

geometry– Damage Tolerance Analysis (DTA) of original & repaired

geometry– Old and new representative flight conditions (RFC):

• Existing vs. improved algorithms to compute spectra – Risk assessment and sensitivity studies– Path forward for continued safety of flight (i.e., “way

ahead”)• In summary, a multidisciplinary, high-priority

effort involving USAF logistics and maintenance engineers teamed with SwRI materials and structural engineers

11

Metallurgical Failure Analysis

• Each crack examined grew in fatigue and initiated (nucleated) at the inside (upper) corner of the LWS radius on the land

• No evidence of a “rogue” flaw was found on either crack• Fracture surface appeared similar for each crack, indicating

similar spectrum loading• Striation spacing measurements were made to estimate stress

ranges that agreed quite well with average stress ranges in the IFF spectrum (T-38’s most severe usage) at this LWS location

• Before becoming a through crack, aspect ratios were 1.83 and 0.86

• Previous DTAs assumed a corner crack initiated (nucleated) at the outer (lower) corner of the land; the DTA used an aspect ratio of 2.0

• The wing crack from TN 64-3261 was very similar to that analyzed as FCL A-9 in previous DTAs

12

Original and RepairedGeometry

• Typical repair is to blend-out crack by removing material from the land• Radius increases from 1.5 to approximately 2.3 inch

Outbd.

Up

A

A

0.313SECTION A-A

Wing LowerSkin

Fwd.

Fwd.

44% Chord Line

W.S. 64.8

0.3350.188

Outbd.

Up

A

A

SECTION A-A

Wing LowerSkin

Fwd.

Fwd.

44% Chord Line W.S. 64.8

0.3

2.3 Radius

Original (with Land) Repaired (Land Milled Out)

13

FEA and DTA ResultsOriginal and Repaired Geometry

Original (Kt = 2.02) Repaired (Kt = 1.84)

• Static Load Case 104: 7.33g symmetric pull-up at 36,000 ft, Mach 1.03• Original NASTRAN -29 wing model validated at FCL A-9 by a full-scale

wing test

14

FEA and DTA ResultsOriginal and Repaired Geometry

• Maximum principal stresses obtained from FEA along the FCL A-9 crack path

• Crack modeled as quarter-elliptical corner crack at edge of plate growing in a tensile stress gradient with transition to a through crack (initial a/c = 1.0)

• AFGROW and tabular Walker model used for 7075-T7351 plate material

• Two IFF stress spectra approaches evaluated

• 2003 T-38 DADTA served as a baseline reference

1

10

100

1000

10000

100000

-15 -10 -5 0 5 10 15 20 25 30 35

Stress (ksi)

Exce

edan

ces

OLD Flight ConditionsNew Flight Conditions

7075-T73 Plateda/dN vs. ∆Κ

1E-08

1E-07

1E-06

1E-05

1E-04

1E-03

1000 10000 100000∆K (psi root in)

da/d

N (i

nche

s/cy

cle)

R=0.1 R=0.5 R=0.8 FIT

TabularWalker Model

FCL A-9 IFF Stress Spectra

15

FEA and DTA ResultsDurability Analysis at FCL A-9

Crack Growth Comparisons

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000

Flight Hours

Cra

ck L

engt

h (in

.)

OLD New TN 64-13261 TN 68-08201 New 0.0082 Initial Flaw

1.5

16

FEA and DTA ResultsIFF Flight Conditions (FCL A-9)

O riginal G eom etry Repaired G eom etryInitial Crack Size (in) 0.05 0.05Critical Crack Size (in) 0.82 0.79Safety Lim it (hours) 414 651In itial Inspection (hours) 207 325Recurring Inspections (hours)

NDI Detectable 0.05 inch 207 325NDI Detectable 0.10 inch 104 205NDI Detectable 0.15 inch 70 120

• Recurring inspection interval used was 125 hours• Recurring interval should be conservative:

• Safety limit calculated to be 1,000 hours (2003)• ½ the safety limit is technically acceptable (450 hour interval)• Field experience with this FCL suggested a “tighter” interval

• 2003 DADTA for FCL A-9 used a NDI “field detectable” flaw size of 0.05 inch (surface eddy current probe)• Initial flaw size also assumed to be 0.05 inch

17

Motivation for Risk Analyses

• Very critical LWS location; short inspection intervals

• Numerous cracks found in field• Numerous assumptions made in DTA of

repair; no two repairs are alike• Toughness dependence on thickness• Cracks initiated (nucleated) on inside

(upper) corner of LWS land and would likely not be detectable with a surface eddy current probe until it was a through crack

18

Development of EIFS Distribution

Data for a PROF Risk Analysis

• Needed to develop an Equivalent Initial Flaw Size (EIFS) distribution

• Deterministic DTA Input:– Stress Intensity Factor – Crack Growth Curve

• Random Variables:– Fracture Toughness– Initial Crack Size Distribution– Maximum Stress per Flight– POD Curve– Repaired Crack Size Distribution

• Aircraft Data:– Number of Locations– Number of Aircraft– Inspection Intervals

t

Kσ

P R O F

POD

a

a

f

a

ap

maxσ K

g

cr

POF

Time

SingleFlight

POF

Time

∆T Interval

InspectionTimes

[Figure taken from PROF Manual]

19


• EIFS is the most significant input for a risk analysis

• No EIFS data exist for FCL A-9• In 1996, SwRI and SA-ALC

documented cracks found at FCL A-9 in nine AT-38B IFF/LIF aircraft (flight hours included)

• These data were combined with the current findings

• Initial flaws were back-calculated: – Original geometry DTA model – Both stress spectra

20


Summary of Defects Found at the Landing Gear Door Radius at WS 64.8 (FCL A-9)

Tail No. Wing No.

A/C Side

Flight Hours

Report Date

Defect Size

(inches) 61-0911 SP 0232 Right 3614.7 10/10/95 0.063 61-3678 SP 0358 Right 3344.8 6/28/95 0.0625 61-0866 SP 0229 --- 3598.9 6/26/95 0.125 61-0891 SP 0264 Right 3542.2 10/11/85 0.19 61-0876 SP 0301 Right 3573.6 9/27/95 0.5 67-14842 --- Right 3458.5 12/2/93 0.125 61-0845 SP 0248 Right 3348.2 10/17/95 0.1 61-0852 --- Right 3464.5 6/10/94 0.43 62-3752 SP 0362 Left 3031.1 5/2/95 0.09 64-13261 SP 0852 Left 3173 4/7/04 0.75 68-08201 SP 0951 Right 2830 3/28/04 1.5

21


Computation of EIFS for In-Service Cracks Found at the Landing Gear Door Radius at WS 64.8 (FCL A-9)

Tail No. Wing No.

Flight Hours

Defect Size

(inches)

EIFS (inches) for Old RFC

EIFS (inches) for New RFC

Empirical CDF

61-0911 SP 0232 3614.7 0.063 0.0070246 0.0042002 0.061 61-3678 SP 0358 3344.8 0.0625 0.0074595 0.0047754 0.149 61-0866 SP 0229 3598.9 0.125 0.0079315 0.0047754 0.237 61-0891 SP 0264 3542.2 0.19 0.0079315 0.0050952 0.325 61-0876 SP 0301 3573.6 0.5 0.0084427 0.0050952 0.412 67-14842 --- 3458.5 0.125 0.0084427 0.0050952 0.500 61-0845 SP 0248 3348.2 0.1 0.0084427 0.0050952 0.588 61-0852 --- 3464.5 0.43 0.008984 0.0054346 0.675 62-3752 SP 0362 3031.1 0.09 0.0095686 0.0061905 0.763 64-13261 SP0 852 3173 0.75 0.0101772 0.0061905 0.851 68-08201 SP 0951 2830 1.5 0.0115883 0.0075436 0.939

22


• Empirical CDF values (for N=11) were computed using Bernard’s approximation for median rank

• The “best fit” to the empirical CDFs were obtained using an Extreme Value distribution:

F(c) = exp[- exp(- (c - m)/α)]where:

c = crack size (inches)m = modeα = scale

Extreme Value Parameter Old RFC New RFC Mode (m) 0.00816 0.005023 Scale (α) 0.00095 0.000640

23


0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1.0

0.000 0.005 0.010 0.015

EIFS (inches)

CD

F

Extreme Value -- Old RFC

Extreme Value -- New RFC

Old RFC - Ranked

New RFC - Ranked

Old Loads

New Loads

EIFS Cumulative Distributions for T-38 FCL A-9 Old and New Representative Flight Conditions (RFC)

24

Risk Analyses and Sensitivity Studies

Other Risk Analysis Input

• Two locations per aircraft• Maximum stress per flight (extreme value

distribution) for each stress spectrum• Fracture toughness (normal distribution)• DTA results (crack growth curves and stress

intensity factor models)• Probability of detection (POD) curves• Note: RFCs (“Old” and “New”) are not two different

spectra, but are two analytical approaches to the use of the IFF spectrum (using the available data)

25


Maximum Stress per Flight DistributionsOld and New Representative Flight Conditions (RFC)

FCL A-9, IFF Old RFC

1.E-03

1.E-02

1.E-01

1.E+00

11 13 15 17 19 21 23 25 27 29 31

Max Stress per 1000 Flight Hours (ksi)

Prob

abili

ty o

f Exc

eeda

nce

Gumbel Parameters: A = 2.30 B = 17.53

FCL A-9, IFF New RFC

1.E-03

1.E-02

1.E-01

1.E+00

11 13 15 17 19 21 23 25 27

Max Stress per 1000 Flight Hours (ksi)

Prob

abili

ty o

f Exc

eeda

nce

Gumbel Parameters: A = 1.43 B = 17.55

26


Probability of Detection (POD) Curves

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0.00 0.05 0.10 0.15 0.20

Crack Size (inches)

Prob

abili

ty o

f Det

ectio

n

Sigma = 0.5

Sigma = 0.75

Sigma = 1.0

Minimum Detectable = 0.0 inchesMedian Detectable = 0.05 inches

Note: Previous DTAs of FCL A-9 used a NDI field-detectable flawsize of 0.05 inches to determine inspection intervals

27


• To start, calculations were performed for a single aircraft to compute the Single Flight Probability of Fracture (SFPOF):– “Original” and “Repaired” (blended) geometry– “Old” and “New” RFCs– Without inspections for 5,000 hours– With recurring inspections at 125 hours– Postulated Inspection Scenario

• Other risk evaluations also considered

28


1E-15

1E-14

1E-13

1E-12

1E-11

1E-10

1E-09

1E-08

1E-07

1E-06

1E-05

1E-04

1E-03

1E-02

1E-01

1E+00

0 500 1,000 1,500 2,000 2,500 3,000 3,500 4,000 4,500 5,000

Flight Hours (IFF)

SFPO

F

Old RFC, Original Geometry

Old RFC, Milled Geometry

New RFC, Original Geometry

New RFC, Milled Geometry

Case: Without Inspections

29


Postulated Inspection Scenario

• The subject aircraft experienced a “mixed” usage prior to being transferred to a new AFB, resulting in approximately 1,450 hours of equivalent IFF usage• To account for the “mixed” usage, the following hypothetical scenario was analyzed:

• 0 hours, begin “equivalent” IFF service• Inspection at 1,000 hours• Inspection at 1,125 hours• Inspection at 1,250 hours• Inspection at 1,375 hours• Inspection at 1,500 hours (and transfer to new AFB)• Inspection at 3,000 hours

30


1E-15

1E-14

1E-13

1E-12

1E-11

1E-10

1E-09

1E-08

1E-07

1E-06

1E-05

1E-04

1E-03

1E-02

1E-01

1E+00

0 500 1,000 1,500 2,000 2,500 3,000

Flight Hours (IFF)

SFPO

F

O ld RFC, Original Geometry

Old RFC, Milled Geometry

New RFC, Original Geometry

New RFC, Milled Geometry

Case: Postulated Inspection Scenario

No Inspections

Inspectat 125 hours

No Inspections

31


Postulated Inspection Scenario

• After 1,500 hours of IFF usage, the SFPOF will rise rapidly to an unacceptable level (“greater than” 1E-07) within the next 500 hours unless inspections are performed (original geometry)

• For the milled-out geometry, this increase in SFPOF occurs after 2,000 hours, and takes 1,000 hours to reach an unacceptable level

32


• Sensitivity studies evaluated various parameters, including:– Mean Fracture Toughness:

• 35 ksi√in (used in 2003 DTA)• 61.1 ksi√in (for thickness of 0.335 in)• 102 ksi√in (obtained in recent tests)• POF criterion is based on toughness• Coefficient of variation (std dev/mean) of 10 percent

– Median Detectable Crack Size:• 0.05 inch (used in DTA)• 0.10 inch• 0.15 inch (crack visible from lower surface by SEC)

33


1.E-15

1.E-14

1.E-13

1.E-12

1.E-11

1.E-10

1.E-09

1.E-08

1.E-07

1.E-06

1.E-05

1.E-04

1.E-03

1.E-02

1.E-01

1.E+00

0 500 1,000 1,500 2,000 2,500 3,000 3,500 4,000 4,500 5,000

Flight Hours (IFF)

SFPO

F

New RFC, Kc = 35

New RFC, Kc = 61

New RFC, Kc = 102

Old RFC, Kc = 35

Old RFC, Kc = 61

Old RFC, Kc = 102

Case: Fracture Toughness - SFPOF for the Original FCL A-9 Geometry w/o Inspections, Both RFCs, and Three Values of Toughness

34


1.E-15

1.E-14

1.E-13

1.E-12

1.E-11

1.E-10

1.E-09

1.E-08

1.E-07

1.E-06

1.E-05

1,000 1,200 1,400 1,600 1,800 2,000 2,200 2,400

Flight Hours (IFF)

SFPO

F

Median Detectable = 0.05 inchesMedian Detectable = 0.10 inchesMedian Detectable = 0.15 inches

Case: Detectable Crack Size - SFPOF for the Original Geometry with 125 hour Recurring Inspections and After 1,000 hour Inspection Free Period for Three

Median Detectable Flaw Sizes and the New RFC

35

Conclusions and Recommendations

• A 125 hour recurring inspection interval at FCL A-9 will maintain risk at an acceptable level

• A fleet-wide assessment of geometry variations at the WS 64.8/44% spar wing radius area should be made

• These risk analyses are preliminary and represent risk for a single FCL and a single aircraft; estimates of fleet-wide risk for different usages can also be made.

• An accurate POD curve for the original and repaired geometry at FCL A-9 is needed

• Currently, PROF cannot account for significantly different post-repair geometries

• Record all field-detected crack lengths and flight hours prior to making repairs

36

USAF Actions and Way Ahead

• Engineering Issues:– A flight loads data recorder program is currently underway to better define

the IFF usage for the T-38 fleet• Additional data will support the baseline Loads/Environment Spectra Survey

(L/ESS) for the T-38 fleet• Data could indicate a more severe IFF spectrum• L/ESS data will then feed the next T-38 DADTA

– In tandem with the L/ESS efforts, a restructured Individual Aircraft Tracking Program (IATP) for the T-38 fleet is also underway, and will assist in monitoring the fleet

• Detect usage changes more proactively• Ability to better evaluate individual aircraft

• NDI Issues:– Better communication between the T-38 ASIP/Engineering/NDI community

and the field-level NDI units is vital, and needs improvement• Program underway for biennial exchanges (site-visits) between Hill AFB

personnel and the T-38 bases• Improved NDI tools (e.g., Rotoscan replacement) and more “user friendly” Tech

Orders now being developed and/or implemented

37

References

1. Cardinal, J.W., Wieland, D.H., and Lankford, J., “Metallurgical and Fracture Mechanics Analysis of T-38 Wing Skin Cracking,” SwRI Project No. 18.08760, USAF Contract No. F42620-00-D-0037, D.O. No. 0048, October 2004.

2. Cardinal, J.W., Wieland, D.H., and Burnside, O. H., “DTA, Risk and Fatigue Analysis for Original and Repaired T-38 Wing Skin Cracks at FCL A-9,” SwRI Project No. 18.08760, USAF Contract No. F42620-00-D-0037, D.O. No. 0048, December 2004.

Lockheed Martin Aeronautics Company

G.R. Bateman, LM AeroP. Christiansen, WR/ALC

28 November 2005

2005 USAF Aircraft Structural Integrity Program Conference

C-130 Center Wing Fatigue Cracking A Risk Management Approach

Lockheed Martin Aeronautics Company 22005 USAF ASIP Conference

Overview

• Background

• Overview of C-130 ASIP � How did we get here?

• Service Cracking - Correlation & Residual Strength Analysis

• A Change in Direction - Risk Assessment Methodology and Results

• Structural Integrity Management Strategies

• Conclusions and Lessons Learned




Background


Background

YC-130 First Flight 23 August 1954

• C-130 has been in continuous production for 50 Years

• 5 Basic Models built & 70 Derivatives Models with a wide range of operational roles

• Operated by more than 60 Nations World Wide− USAF operates nearly ½ of all C-130/382 Models in service

C-130 Major Model Production

231230

135

5311165

C-130AC-130BC-130EC-130HC-130J


BackgroundC-130 Center Wing Box


Background

Center Wing Section View (Typical)



Overview of C-130 ASIPHow Did We Get Here?



• The C-130A (1955) was designed prior to requirements for service life expectancy, fatigue endurance, damage tolerance, and durability testing

• The original C-130E (1961 � 1968) Center Wings experienced significant fatigue cracking after only 6 years of service

• C-130 Center Wing now in service was designed in 1968:− Retrofitted to all C-130B and Early C-130E aircraft− Fatigue Tested to 43,000 FH of E Usage with few cracks

• C-130H Wing Durability Test conducted 1988 to 1992:− Determine Service Life of Outer Wing & Assess impact on increased

usage severity on the Center Wing− 60,000 FHs of MAC Usage (Twice as severe as DTA Baseline Usage)



• Wing Durability Test (1992) results used to predict the time to reach generalized cracking

• Test Spectra Max Load approx 60% DLL, insufficient to cause failure during test

• MSD/MED caused Failure of Center Wing Lower Surface at 100% DLL, Residual Strength Test at 120,000 EBH

• 45,000 EBH established as upper bound Economic Service Life (Includes a Scatter Factor of 2)

C-130H Wing Durability Test Results

0

100

200

300

400

500

600

700

800

900

1000

0 20 40 60 80 100 120 140

Equivalent Baseline Hours (X 1000)

Cum

ulat

ive

# of

Cra

cks

Center Wing Service Life

END of B/EWING TEST

CW Test Loading Reduced

Residual Strength

Test

Center WingOuter Wing



Service Cracking Correlation Analysis



• 1995 - Cracking Rate Projections in Service Life Analysis using Fatigue (Stress-Life) analysis methodology:− Initial Assessment over-estimated severity of C-130H Wing Durability

Test− Center Wing Service Life (economic) estimated at 60,000 EBH

• 2002 - Increased rate of service crack discovery:− More than expected on C-130E Training Aircraft− Less than expected on Special Operations C-130H derivative Aircraft

• 2004/2005 - Updated Operational Loads Programs incorporated into IATP corrected discrepancy in relative severity between different usage groups

• 2005 � Crack Growth Methodology replaced Fatigue Methodology in determining Service Life



Lower Surface FWD Panel

WS 215

Center to Outer Wing Joint Access Bolt Cut-outs

Fatigue Cracks

MSD in Center Wing Lower Surface Skin Panel



FWD

Beam Cap crack (severed)

Skin Panel crack

MED in Center Wing Lower Surface



Front Beam Web

Web Splice StiffenerLower Surface Panel

Stringer

FWD

TL 100

H 1446

Front Beam Cap

MED - Simultaneous Cracks in Center Wing Lower Surface

Beam Cap and Skin Panel



MSD and MED In Center Wing Lower Surface Panel



• Correlation of Analysis to Test and Service Cracks

• Equivalent Baseline Hours (EBH) determined from updated IATP

• No clustering of MDS or Usage type could be determined, with the exception of a weather reconnaissance aircraft

• Service Cracks regressed to define EIFS Distribution

Discovered Test and Service Cracks

REPORTED TEST AND SERVICE CRACKSCENTER WING LOWER SURFACE PANEL AT WS 61

0

1

2

3

4

5

6

Equivalent Baseline Hours (EBH)

Cra

ck L

engt

h (in

ches

)

Safety LimitDurability LimitMilitary Service CracksCommercial Service CracksAircraft Inspected - No CracksTest Cracks

a

a


• Service Cracks regressed to a common crack size (a = 0.20 in, a high POD for ABHEC)

• Reliability Analysis excluding “survivors”, the aircraft inspected with no cracks detected

• Discovery of Service Cracks fall within a log-normal distribution of EBH

• Log-Normal (green line) used to derive EIFS


Distribution of Time (EBH) to Common Crack Size

Distribution of Time to Common Crack SizeLower Surface Panel at WS 61

0

10

20

30

40

50

60

70

80

90

100


Cum

ulat

ive

Prob

abili

ty (%

)

Reliability Results of RegressedService Cracks

Fitted Log-Normal Distribution

Log-Normal Dist. FittingMean: 36,714 EBH

Std. Dev. 0.115



• EIFS Distribution calculated two ways:− From Log-Normal

Curve Fit to time to 0.20 inch crack

− Weibul Distribution directly from regressed service cracks

• Both methods give a mean EIFS of approx 0.005 in.

• Log-Normal method results in larger EIFS at top of distribution

Equivalent Initial Flaw Size (EIFS) Distribution Lower Surface Panel at WS 61

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1.0

0.0000 0.0025 0.0050 0.0075 0.0100 0.0125 0.0150 0.0175 0.0200EIFS (in.)

Cum

ulat

ive

Den

sity

Fun

ctio

n

EIFS - Derived from Log Normal Dist'b of 0.2 in cracks

EIFS - Derived from Weibul Distb'n of Regressed Cracks

EIFS of Regressed Service Cracks

50% Probability of a 0.0047 in. EIFS (Log Norm Regression of 0.20 in. Service Cracks)

50% Prob of a 0.0052" EIFS (Weibull Fit of Regressed Service Cracks)

Equivalent Initial Flaw Size (EIFS) Distribution



Service Cracking Residual Strength Analysis



• Residual Strength Analysis of Service Cracking scenarios conducted in areas difficult to inspect:− Primary (lead) cracks− Severed Single Element (Discrete Source Damage)− MSD & MED Scenarios

• Constructed a Detailed FEM of Center Wing Lower Surface:− Skin Panel, Stringers, Beam Caps and fasteners modeled

separately− Embedded into Full Airframe FEM− Correlated to available Strain Survey Data



• RH Center Wing Detail FEM incorporated into airframe level FEM

• Stresses correlated to measured strain data

• Used for Residual Strength Analysis and crack geometric restraint solutions

Center Wing Detailed FEM

LM Aero C-130 Airframe Level FEM



Skin

Ribs

Front BeamStringers

Rear Beam

WS 61

WS 22

0

Center Wing Detailed FEM

WS 0



RIBS STRINGERS26,000 elementsFastened to lower surface skin

LOWER SURFACE SKIN14,000 Elements

Detail Mesh at WS 178

Detail Mesh at WS 61

Front and rear beam lower capsfastened to lower surface skin. Upper caps are fused to upper skin.

FRONT BEAM

REAR BEAM

Center Wing Detail FEM Exploded View



• FEM Stresses Correlated to Measured Stresses on Wing Durability Pre-Test Survey

• 90% of locations within ± 10% of Measured Stress

Wing Durability Validation UpbendingPreTest Strain Survey

-16000

-12000

-8000

-4000

0

4000

8000

12000

16000

-16000 -12000 -8000 -4000 0 4000 8000 12000 16000

Measured Stress, psi

Pred

icte

d St

ress

, psi

Center Wing Lower Surface FEM Stress Validation Wing Durability PreTest Strain Survey

0.1

1.0

10.0

Probability of Prediction Ratio

Rat

io o

f Pre

dict

ed S

tress

to

Mea

sure

d St

ress

0.01 0.1 1 10 30 7050 90 99 99.9 99.99

Predicted stresses are within +/- 10% of measured values

Center Wing Lower Surface FEM Stress Correlation



Lower Surface Skin Panel Stress Distribution

►O

utbo

ard

►Aft

WS 140

WS 0

Lower Surface Skin; FS 517- FS 597

Y-Component

AFT

Intact Panel

Severed Forward Skin Panel15% Increase in Mid Panel

average stress



Lower Surface Stringer Stress Distribution

►O

utbo

ard

►Aft

WS 140

WS 0

Stringers 12 thru 24

Y-Component

Intact Panel

Severed Forward Skin Panel350% Increase in Stringer foot

average stress (across panel crack)



Stringer 13; Aft flange

FEM with �no hole�

12 13 14 15 16

FEM with �detailed hole�

Lower Surface Forward Panel Severed

Stringers Shown (Panel Removed)

OUT’BD

WS 60

Lower Surface Stringer Foot Stresses � Detailed Fastener Hole



• Residual Strength Analysis Conclusions:− Skin panel can tolerate a long crack under normal operational

loads− Lower Surface structure is Fail-Safe with a single panel

severed (providing there is NO adjacent panel or stringer cracking)

− Stringers cannot tolerate ligament cracks with a severed skin panel at high loads (high ZFW & load factors above 2g)

− With the likely presence of adjacent stringer cracking as the center wing approaches the Service Life, Inspections must detect panel cracking prior to fracture to meet Damage Tolerance Requirements



A Change in DirectionRisk Assessment Methodology and Results



• USAF ASIP Independent Review Team (IRT) provided direction to perform Risk Assessments of:− Discrete Source Damage scenario of a severed skin panel− Intact structure, Principal Structural Element primary fatigue

crack propagating undetected in the skin panel

• LM Aero Crack Growth Methodology used throughout:− Equivalent Baseline Hours (EBH) determined from IATP− EIFS Distribution from Service Crack Correlation Analysis− Residual Strength Analysis from Center Wing Detail FEM− Several Usage Groups considered for Load (Stress)

Occurrences

• Single Flight Probability of Failure (SFPoF) vs EBH used to quantify the Risk



Stress Occurrences

Residual Strength

EIFS Distributions

Crack Growth vs EBH

Single Flight

Probability of FailureSingle Flight Probability of Failure (SFPoF)

Center Wing Lower Surface

1.E-09

1.E-08

1.E-07

1.E-06

1.E-05

1.E-04

1.E-03

1.E-02

1.E-01

1.E+00


SFPo

FSevere Military Usage

Typical Military Usage

Typical Commercial Usage

Cumulative Max Stress Occurrences Per Flight Lower Surface Panel at WS 61

1E-10

1E-09

1E-08

1E-07

1E-06

1E-05

1E-04

1E-03

1E-02

1E-01

1E+00

1E+01

1E+02

Max Stress (Ksi)

Cum

ulat

ive

Occ

urre

nces

per

Flig

ht

Severe Military Usage Typical Military UsageTypical Commercial Usage

Equivalent Initial Flaw Size (EIFS) DistributionLower Surface Panel at WS 61

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1.0

EIFS

Cum

ulat

ive

Den

sity

Fun

ctio

n

Crack Growth vs EBHLower Surface Panel at WS 61

0

2

4

6

8

10

12

14

16

18

20


Cra

ck L

engt

h a

(in)

Residual StrengthLower Surface Panel at WS 61

0

10

20

30

40

50

60

70

Crack Length

Stre

ss (K

si)



• Comparison of LM Aero C-130 Risk results to PROF

• LM Aero probability of Stress occurrence is slightly higher than Gumbel representation

• Results nearly identical when Gumbel distribution is used in LM Aero C-130 Risk program

Single Flight Probability of Failure (SFPoF)Center Wing Lower Surface - Single Zone

1.E-09

1.E-08

1.E-07

1.E-06

1.E-05

1.E-04

1.E-03

1.E-02

1.E-01

1.E+00


SFPo

F

LM Aero Risk Analysis

PROF

LM Aero Risk with Gumbel Stress Distribution



• Single Flight Probability of Failure for the Center Wing Lower Surface for “Intact” structure

• At low EBH values, Risk is Usage Dependent, set by occurrence rate of high maneuver loads

• At high EBH values, Risk is independent of usage, set by occurrence rate of typical gust loads

Single Flight Probability of Failure (SFPoF)Center Wing Lower Surface

1.E-09

1.E-08

1.E-07

1.E-06

1.E-05

1.E-04

1.E-03

1.E-02

1.E-01

1.E+00


SFPo

F

Severe Military Usage

Typical Military Usage

Typical Commercial Usage

Unacceptable

OK

RiskMitigation



• �Discrete Source Damage� Analysis indicated unacceptable levels of Risk exist in the event of a single lower surface skinpanel failure for Severe Military Usage:− Stringer attachment flanges experience significant yielding at

loads approaching 80% of Design Limit− Occurrence Rate of high maneuver loads approaching 80%

DLL exceeds 1 x 10-4

− Conclusion is that Risk Management Strategy must eitherdetect panel cracking prior to fracture; or remove center wing from service when risk exceeds an acceptable level



Structural Integrity Risk Management Strategies


Structural IntegrityRisk Management Strategies

• Numerous Risk Mitigation Strategies have been employed by the USAF ASIP Manager:− Operational Flight Restrictions Imposed on USAF aircraft to

reduce maximum wing up-bending load to below 60% of Design Limit at 38,000 EBH

− Many TCTOs released to inspect for localized fatigue cracking− TCTO released to inspect for generalized cracking of Lower

Surface of selected Center Wings− Established an upper bound Service Life of 45,000 EBH,

grounding of high time C-130 aircraft



• For non-USAF operators, LM Aero has released two major Service Bulletins:− 82-788/382-57-84 Operational Usage Evaluation and Service

Life Assessment• Requires all operators to perform a usage assessment to

determine EBH on the Center and Outer Wings• Recommends an operational usage evaluation, flight restrictions,

inspections and grounding depending on EBH Status− 82-790/382-57-85 Lower Surface Generalized Cracking and

Widespread Fatigue Damage Inspection Requirements:• Provides inspection instructions to detect generalized cracking

and/or possible onset of WFD• Covers nearly 100% inspection of center wing lower surface

panels, stringers and beam caps



• Evaluation of Risk Mitigation through Inspection in progress:− Highly dependent on Probability of Detection (POD) and

Probability of Inspection (POI), require more data− Consider effects of fastener oversize and cold working− Intrusive inspections are likely to find more damage

necessitating repair, economic factors may drive eventual replacement



Conclusions and Lessons Learned


Conclusions and Lessons Learned

• Single Panel Failure must be prevented either by a combination of restrictions and inspection, or removal from service prior to an undetected fatigue crack reaching critical size

• Initial levels of risk out to the �rogue� flaw safety limit depend largely on the occurrence rate of high maneuver loads

• Point at which Risk increases rapidly largely depends on the EIFS distribution

• Restrictions are not effective once unacceptable level of Risk is reached (i.e. failure can occur from typical gust loads)

• Service Cracking Data is vital to the Risk analysis process

1

2005 USAF ASIP Conference

The C-17 Aircraft External Load to Local

Stress Spectrum Transformation with Correlated Results

NOV 2005Archie Woods - USAF, C-17 ASIP Manager

Ko-Wei Liu � The Boeing CompanyRick Selder � The Boeing Company

2


C-17A Globemaster III

3


Outline

● Acknowledgements● Decision to Assess Local Stress Spectrum● Basis for the C-17 Weapon System Approach● Stress Spectrum Approach for the C-17 Aircraft● Wing Local Stress Spectrum Development● Strain Survey Overview● Wing Local Stress Correlation Results● Observations● Future Plans

4


Acknowledgements

Ko-Wei Liu, The Boeing Company, Technical Specialist

Rick Selder, The Boeing Company, Spectrum Development

Dr Joseph Gallagher, USAF, USAF ASIP Manager

Hugo Guzman, The Boeing Company, DADTA Manager

5


Decision to Assess Local Stress Spectrum

● Identify high cost or significant economic impacts to the United States Air Force

● Improvements to stress spectrum development procedures or methodology will focus on major structural components where predicted crack growth is greatest for actual usage

● Concentrate on primary structure within major components

● Identify potentially difficult/intrusive inspection areas

Wing major structure highest priority for local stress spectrum correlation assessment – Focus of current review

Future local stress spectrum correlation assessments will occur in the following order: Fuselage-to-Wing Joint,

Empennage, Engine Pylon, and Fuselage

6


● The USAF C-17 Program Office and The Boeing Company decided it was not expeditious/economical to perform a Finite Element Model (FEM) analysis for every condition to directly calculate local stresses

● Fuselage, wing, horizontal stabilizer, engine pylon, and vertical stabilizer collectively have over 200,000 FEM elements.

Basis for C-17 Weapon System Approach

● C-17 aircraft airframe airworthiness certification required a large number of loading conditions to define the operational loading environment

7


Basis for C-17 Weapon System Approach

FEM Statistics

1K2K1KEngine Pylon

25K30K40KVertical Stabilizer

11K35K20KHorizontal Stabilizer

42K140K73KWing

45K150k77KFuselage

NodesDegrees

of Freedom

ElementsMajor Structure

Note: FEM represents LHS of aircraft only

8


● Define a matrix of external load cases that best represent the range of loads expected during normal operations� Minimum of 6 external load cases; one for each load axis� Additional cases improve confidence and error checking

● Perform a FEM analysis for each external load case of the matrix; analysis outputs local stresses

● Perform a regression analysis to calculate external load to stress transfer coefficients for specified structural location� Selected structural locations resulted in the ability to

calculate local stresses for numerous control points● Calibrate regression equations or FEM using static

and/or strain survey of full scale test article● For desired loading condition, apply stress transfer

coefficients to external loads to obtain local stresses for control points

Stress Spectrum Approach for the C-17 Aircraft

9


Wing Local Stress Spectrum Development

● Description� Wing assembly is a continuous monocoque construction

with 6 attach locations on top of the fuselage; for drag & vertical loads

� Fuel may be loaded in the entire wing span● External loads were distributed on wing FEM for a total

of 12 maneuver and ground external load cases� Fuselage high/mid/low gross weight (GW) with

high/low/low fuel weight at 1.0g vertical load factor, and cruise speed and altitude

� Design limit cases for low/high GW with low/high fuel weight from 2.0g to 3.25g vertical load factor at low speed or cruise speed

� Ground & taxi for high/mid/low GW with high/mid/low fuel weight

● An additional pressure-only external load case is applied to the wing; differential cabin pressure at 7.8 psi on wing center

10


Wing Local Stress Spectrum Development

TCSvCMvC TSvMv ++=σ

PCShCMhC PShMh +++

● FEM stress for each case and associated external elastic axis loads were input into a least squares regression analysis to obtain stress transfer coefficients for a selected control point

● For the same control point, the FEM stress due to cabin pressure (P) was divided by 7.8 psi to obtain stress due to unit cabin pressure� Linear, proportional relationship

● Using the following equation the stress for any given loading condition was calculated:

11


Strain Survey Overview

● Full Scale Durability (FSD) test article, D1, was used for strain gage correlation of the local stress transformation equations� Four (4) flight conditions

� Symmetric maneuver � wing up bending� Symmetric taxi � wing down bending� Gust, low altitude cruise � heavy cargo� Gust, low altitude cruise � light cargo

● Conditions chosen for the ability to provide good predictions at maximum/minimum of stress range

● Identified actuator contributions to strain gage loads� Actuator loads were transformed to an external load

reference axis, summed, and then used in the local stress transformation equations

● Internal pressure effects are summed when appropriate

12


Wing Local Stress Correlation Results

Wing Lower Skin Panel - Maneuver (up-bending)

σ(gage), ksi

σ(predicted), ksi

15.81617.200

15.44115.862

18.85119.534

RightWing

Left Wing

13

2005 USAF ASIP Conference Wing Local Stress Correlation Results-ContinuedWing Lower Skin Panel � Left and Right● Predicted versus Measured Stress

Wing Lower Skin Panel

-20

-15

-10

-5

0

5

10

15

20

25

-20 -15 -10 -5 0 5 10 15 20 25Measured Stress (ksi)

Pred

icte

d St

ress

(ksi

)

Over Prediction

14


Wing Local Stress Correlation Results-ContinuedWing Lower Skin Panel � Left and Right● Predicted versus Measured Stress

0.1 1 10

Predicted/Measured

Prob

abili

ty o

f Occ

urre

nce

C-17 Lower Wing Surface predicted / measured - stressLinear (C-17 Lower Wing Surface predicted / measured - stress)

99.99%

99.9%

99%

90%

50%

10%

1%

0.1%

0.01%

Range: 10% over predict to 10% under predict

15


Wing Local Stress Correlation Results-Continued

Wing Upper Skin Panel - Taxi (down-bending)

σ(gage), ksi

σ(predicted), ksi RightWing

Left Wing

13.19412.899

14.66416.574

16



Wing Upper Skin Panel � Left and Right● Predicted versus Measured Stress

Wing Upper Skin Panel

-30

-25

-20

-15

-10

-5

0

5

10

15

20

25

-30 -25 -20 -15 -10 -5 0 5 10 15 20Measured Stress (ksi)

Pred

icte

d St

ress

(ksi

Over Prediction

(ksi

)

17



Wing Upper Skin Panel � Left and Right● Predicted versus Measured Stress

0.1 1 10

Predicted/Measured

Prob

abili

ty o

f Occ

urre

nce

C-17 Upper Wing Surface predicted / measured - stressLinear (C-17 Upper Wing Surface predicted / measured - stress)

99.99%

99.9%

99%

90%

50%

10%

1%

0.1%

0.01%

Range: 10% over predict to 10% under predict

18


Observations

● Engineering tools and methodology used for the C-17 wing local stress transformations provide good correlation� Probability of local stress predictions being less than the

actual stress will only slightly occur for a magnitude less than a reasonable percentage of 10% from the exact value

� A portion of the inaccuracy may be inherent in stresses obtained from the FEM

● Correlation using probabilities provides an additional description for accuracy of the engineering prediction techniques and tools

19


● Assess the need to improve wing predictions� Economic and safety impact to the USAF due to under-

prediction of damage; the extent of actual accumulated damage between inspections may require a more extensive repair or modification

� Inspection burden to aircraft user is a consideration; important to optimize maintenance while maintaining mission capability

● Assessments will continue on the other strain gage correlations using the probabilistic approach; Fuselage-to-Wing Joint, Empennage, Engine Pylon, and Fuselage

● All Finite Element Models (FEM) are candidates for revision � Improve the idealization of the structures and correct any

possible anomalies found over the years● DADTA analysis will be updated in the future using the

refined stresses

Future Plans

Air Force Institute of Technology, Warsaw, PolandAir Force Institute of Technology, Warsaw, Polandwww.itwl.pl

Capt. Krzysztof Dragan, MSc Eng., AFIT, Warsaw, Poland

Lt.Col. Sławomir Klimaszewski, PhD Eng., AFIT, Warsaw, Poland

Maj. Andrzej Leski, PhD Eng., AFIT, Warsaw, Poland

NDE APPROACH FOR STRUCTURAL INTEGRITY EVALUATION OF

MIG-29 AIRCRAFT

- ASIP 2005 – Memphis-


• Introduction

• Fleet Status

• NDE as a tool for problem analyzing

• Our concerns

• Methods of inspection

• Examples and results & Encountered problems

• Future plans

• Summary/Questions

Presentation Outline:



MiG-29 Fulcrum

Technical data:

Maximum weight: 17 700 kg

Maximum range: 2200 km (*)

Climb velocity: 330 m/s

Maximum speed: 2230 km/h

Versions:

- MiG-29 (Combat)

- MiG-29 UB (Combat - training)



Fleet Status


Polish MiG-29 Fleet

”Polish” ”ex-Czech” ”ex-German”

• Generally Safe Life

• Few ”Damage Tolerance” elements

• Damage Index (based on nz) for ”ex - German” a/c : 1,3÷1,9

• Need for On – Condition – Maintenance (OCM)


Non Destructive Evaluation

NDE – a process used to detect and characterize anomalies in materials and structures

Problem Appropriate Technique

Data Acquisition

Signal Processing

Feature Classification

Analysis/Solving Input Parameters Raw Data Measured Value

ORFeature Vector

ClassificationDecision

Why NDE ?

• detect and characterize defects

• ensure reliability

• extend the service life



Our Main Concerns

• Fastener structures (surface & subsurface cracks, hidden corrosion)

• Spot welded structures (surface cracks, corrosion);

• Composite structures (delaminations, disbonds, porosity, inclusions)

• Honeycomb structures (disbonds, water ingress)



Methods of inspection (1)

---+b) Hidden Corrosion

-+++Composite structures

--++Honeycomb structures

ECUTShearographyD-SightMethod

+---Spot welded structures

++--a) Fatigue cracks

Fastener structures:

Damage



Methods of Inspection (2)

Advantages of automated systems:

• Automated Inspection (fast, reliable)

• Human Factor reduction

• Computer aided analysis and data collection

• Data storing and manipulating

• Data comparison



Examples & Results (1) – Hidden Corrosion

DAISTM (DSightTM Aircraft Inspection System)

•At Present DAIS I & DAIS II •Based on Double Pass Retroreflection; •D-SightTM effect converts local surface curvature to gray scale changes;• Detection of hidden corrosion (visible by pillowing) in horizontal and vertical (lap splices) - DAIS 250C (250 Cx)• At present it is used by AFIT in:

! MiG-29, Su-22, Jak-40 a/c;! Mi-8, Mi-17 and Mi-14 inspections;


DAIS 250C


Examples & Results (1a) - Basic Principles

• Flat surface causes reflection of uniform light intensity distribution

• Local distortions cause local disturbances in light intensity reflection (especially around fasteners)



Examples & Results (1c)-Data AnalysisPlanar Structure – Data presented with colors

• Quick structure analysis;

• Easy to interpret;

• Different damages marked with different colors;

• Finding exact damage position;


cburns

click on photo to view video clip


• Flexible tracks system• High speed inspection of large areas• Vacuum pump system enables vertical and bottom wing skin inspection

• Methods of inspection:! Ultrasound (longitudinal & shear wave, P-E, T-T)! Eddy Current (single & dual frequency)! Resonance (H-F, Pitch –Catch, MIA)! Multiple methods inspection

• A-scan, B-scan i C-scan

Mobile AUtomated Scanner - MAUS


Systems used for advanced inspection (1)


Systems used for advanced inspection (2)

Main unitPersonal viewing system

Video recorder

Scanning Sensor 303

Scanning Sensor 307

MOI – Magneto Optic (Eddy Current) Imaging (System)

• Very versatile and efficient

•Based on Eddy Current technique

•Fast and reliable

•Data recording

•Basic operator skills

• Limited to depths < 4 mm



Examples & Results (2) - Fatigue CracksMiG-29 wing skin eddy current testing:

• Skin Crack;

• Rib Damage

Skin crack Rib damage

MAUS IV/V System Wing Skin Inspection



Examples & Results (3) – Cracks in fastener structures

Cracks in top skin

Ultrasound Angle Beam Inspection with the use of Through Transmission and Pulse Echo technique

4 channel TT 1 channel PE



Examples & Results (2a) – Cracks in fastener structures

Eddy Current MAUS V Inspection Results



Examples & Results (2a) – Cracks in fastener structures

Magneto-Optic Imaging Technology:• very fast;• fairly simple in use.

Eddy Current – C-scanA

A

A

B

B

BC

C

C


cburns

click on icon at right to view video clip


Examples & Results (3) – Cracks in multilayer structures

Crack in top skin

Crack and hole

in bottom

skin

Eddy current inspection, total

thickness –3mm



Examples & Results (4) - Cracks in thick structuresUsing ultrasound and rotating scanner:

• Four channel inspection

• Fixed angle setup for depth range

• Exact crack location

• Rotation to find crack with different orientation



Examples & Results (5) - Cracks in fastener holesUsing BHEC:

• Necessity to remove fasteners

• Fast and reliable

• Exact crack location

• Different diameters

• Easy to inspect



Examples & Results (6) – UT&EC Wing Skin Inspection

UT Amplitude Wing Skin Inspection

EC HF Wing Skin Inspection

• Thickness change monitoring;

• Crack Inspection

• UT A & EC C scan data presentation



Examples & Results (7) - Spot Weld Inspection

• Crack & Corrosion Inspection

• MOI much more efficient for crack inspection

• Spot Welds indications very similar to crack tip indication *

* Magneto Optic Imaging Technology: A New Tool for Aircraft Inspection�, AMPTIAC, QUARTERLY Volume 6, Number 3



Examples & Results (7a) - Spot Weld Inspection

Spot Weld Indications- ASIP 2005 – Memphis-


Examples & Results (8) - Vertical Stabilizer UT Inspection

MAUS V inspection

• CFRP skin inspection

• UT PE single sensor inspection

• Damages detection:

! Delamination

! Disbond

! Inclusion

! Porosity


cburns

click on photo to view video clip


Examples & Results (8a) - Vertical Stabilizer UT Inspection

Amplitude C scan data TOF C scan data



Examples & Results (8b) - Vertical Stabilizer UT Inspection

•Delamination – longeron structure

• Smaller around fasteners



Examples & Results (8c) - Vertical Stabilizer UT Inspection

• Foreign object inclusion



Examples & Results (8d) - Vertical Stabilizer UT Inspection

UT C-scan: TOF Mode UT C-scan: Amplitude Mode

Porosity



Examples & Results (8e) - Vertical Stabilizer UT Inspection

Disbonds



Examples & Results (8f) - Vertical Stabilizer UT Inspection – time comparison

44Analyzing

3264Inspection

MAUS VMAUS IV


Future NDE plans (1):

Inspection for damages like:

• disbonds (skin to honeycomb)

• delaminations (composite)

Methods of inspection:

• Low Frequency acoustic:

! MIA, Pitch – Catch, Resonance MIA technique:

• skin to honeycomb disbond



Shearography Inspection:Inspection honeycomb & composite structures:

• disbonds (skin to honeycomb)

• delaminations (composite)

• honeycomb damages

• water ingress


1

2

(1) Flaws different sizes with water inclusions

(2) Shearography with vacuum & thermal loading


Conclusions / Summary

• All inspections were In Service Inspections

• Described NDI techniques are in use during inspections and:

• Are increasing measuring possibilities

• Are less time consuming and allow for reliable inspection

• Enable „Human Factor” reduction

• Make it possible to apply NDE for new problems solving

• Enable storing and analyzing data as well as over time monitoring

• Gained experience allow for NDI usage to our other aging aircraft programs

• Next generation a/c inspections: CASA-295 M (new), F-16 (new), C-130 (aged)



Questions ?



Thank you for your attention !

[email protected]


Damage Tolerance Risk Assessment of T-38 Wing Skin Cracks...• T-38 NDI team of experts assembled...

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Transcript of Damage Tolerance Risk Assessment of T-38 Wing Skin Cracks...• T-38 NDI team of experts assembled...