replica-based crack inspection

REPLICAREPLICA--BASED CRACK BASED CRACK INSPECTIONINSPECTION

John A. NewmanJohn A. NewmanU.S. Army Research LaboratoryU.S. Army Research Laboratory

Scott A. WillardScott A. WillardLockheed Martin Space OperationsLockheed Martin Space Operations

Stephen W. Smith and Robert S. Stephen W. Smith and Robert S. PiascikPiascikNASA Langley Research CenterNASA Langley Research Center

INTRODUCTIONINTRODUCTIONCracks found in Space Shuttle Main Cracks found in Space Shuttle Main Engine LHEngine LH22 feedlinefeedline flowlinersflowliners (2002)(2002)–– Ranged from 0.1 inch to 0.6 inch longRanged from 0.1 inch to 0.6 inch long–– Weld repaired, polished, and returned to Weld repaired, polished, and returned to

flightflight–– NDE: No cracks >0.075 inches long existNDE: No cracks >0.075 inches long exist

Revisited in 2004Revisited in 2004–– Unable to ensure safe flight with 0.075Unable to ensure safe flight with 0.075--

inchinch--long cracklong crack

FLOWLINER DESCRIPTIONFLOWLINER DESCRIPTION

Orbiter aft

Engine cavity LH2 feedline

FLOWLINER DESCRIPTIONFLOWLINER DESCRIPTIONLHLH22 consumptionconsumption–– 385,000 gallons385,000 gallons–– 8.5 minutes 8.5 minutes –– Each engine consumes Each engine consumes

15,000 gal/min15,000 gal/min–– Flow induced stress Flow induced stress

cycles in kHz rangecycles in kHz range–– Millions of stress Millions of stress

cycles per flightcycles per flight–– Opportunity for FCGOpportunity for FCG

Release of debris is a Release of debris is a safety concernsafety concern

LH2 flow

US flowliner

Bellows

DS flowliner

PROBLEMPROBLEM

Analysis: Unsafe conditions may occur Analysis: Unsafe conditions may occur for cracks > 0.005 inch longfor cracks > 0.005 inch long

Improved eddy current unable to detect Improved eddy current unable to detect 0.0050.005--inchinch--long crackslong cracks

Must develop an NDE method capable Must develop an NDE method capable of finding 0.005of finding 0.005--inchinch--long crackslong cracks

PROPOSED SOLUTIONPROPOSED SOLUTIONUse surface replicas as an NDE methodUse surface replicas as an NDE methodSurface replicas used for decades to Surface replicas used for decades to monitor small cracks (<0.005 inch)monitor small cracks (<0.005 inch)RecentlyRecently--developed siliconedeveloped silicone--based based replicas better suited for inspectionreplicas better suited for inspection

Acetate tape replica Silicone-based replica dispenser

EXPERIMENTAL PLANEXPERIMENTAL PLAN

Feasibility study to evaluate Feasibility study to evaluate ability of siliconeability of silicone--based based replicareplica–– Generate fatigue cracks in Generate fatigue cracks in

laboratory specimenslaboratory specimens–– Compare crack lengths from Compare crack lengths from

SiliconeSilicone--based replicas (no load)based replicas (no load)AcetateAcetate--tape replicas (max load)tape replicas (max load)Direct examination (no load)Direct examination (no load)

FATIGUE TESTINGFATIGUE TESTINGSpecimens used to Specimens used to simulate simulate flowlinerflowliner slot slot geometry and stressesgeometry and stressesTesting interrupted Testing interrupted periodically for slot periodically for slot surface replicationsurface replicationObjective: produce a Objective: produce a distribution of cracks distribution of cracks for replica studyfor replica study

REPLICA ANALYSISREPLICA ANALYSISReplica preparationReplica preparation–– Sectioned in 4 piecesSectioned in 4 pieces–– Grounded on metallic slideGrounded on metallic slide–– Coated with metallic materialCoated with metallic material

Examined in an SEMExamined in an SEMInitial scan at 50Initial scan at 50--100X100X–– Surface finish, scratches, etc.Surface finish, scratches, etc.

Crack scan at 400Crack scan at 400--700X700XOrientation markings

EXPERIMENTAL RESULTSEXPERIMENTAL RESULTSCrack found after Crack found after 50,000 cycles50,000 cycles–– Surface crackSurface crack–– 0.008 inches long0.008 inches long

100 µm

10 µm

CRACK LENGTH COMPARISONCRACK LENGTH COMPARISON

50 µm50 µm

Acetate replica (loaded) – 163 µm Silicone replica (no load) – 199 µm

Specimen (no load) – 194 µm

50 µm

EXPERIMENTAL RESULTSEXPERIMENTAL RESULTS

Crack #1 – 0.012”

200 µm

Crack #3 – 0.001”

Crack #2 – 0.004”

100 µm

3 cracks found after 3 cracks found after 50,000 cycles50,000 cycles–– 2 surface cracks2 surface cracks–– 1 corner crack1 corner crack

CRACK LENGTH COMPARISONCRACK LENGTH COMPARISON(Crack #2)(Crack #2)

20 µm 20 µm



20 µm

CRACK LENGTH COMPARISONCRACK LENGTH COMPARISON(Crack #3)(Crack #3)

10 µm 10 µm



10 µm

CRACK DETECTION FINDINGSCRACK DETECTION FINDINGS

SiliconeSilicone--based replicas:based replicas:–– Able to find cracks shorter than 0.005 inchAble to find cracks shorter than 0.005 inch–– More accurate crack lengths than acetate More accurate crack lengths than acetate

tape replicastape replicas–– More convenient than acetate tape replicasMore convenient than acetate tape replicasNow need to show:Now need to show:–– PostPost--repair crack detection capabilityrepair crack detection capability–– ReproducibilityReproducibility–– No debrisNo debris

CRACK DETECTION AFTER CRACK DETECTION AFTER POLISHINGPOLISHING

FlowlinerFlowliner slots were polished after 2002 slots were polished after 2002 crack repaircrack repair–– Removal of small cracks and punch Removal of small cracks and punch

damage from manufacturedamage from manufactureOne orbiter has not flown since One orbiter has not flown since flowlinerflowliner slot polishingslot polishingConcern about postConcern about post--polishing crack polishing crack detectiondetection–– Crack mouth potentially filled with materialCrack mouth potentially filled with material

POLISHED CRACK DETECTIONPOLISHED CRACK DETECTION

200 µm

50 µm

200 µm 50 µm

200 µm

50 µm

Initial crack

After polishing

After polishing + 1 load cycle

SURFACE FINISH QUALITYSURFACE FINISH QUALITYCrack #1

Crack #2

Crack #3

Crack #4

Crack #5

Crack #6

Crack #7

PitPit--like damage from like damage from punching not punching not completely removed completely removed by polishingby polishingAt least 7 fatigue At least 7 fatigue cracks initiated by cracks initiated by 50,000 cycles50,000 cyclesQuality of surface Quality of surface finish is importantfinish is important

200 µm

OTHER TYPES OF DAMAGEOTHER TYPES OF DAMAGE

200 µm200 µm

200 µm 200 µm

Pit damage Tool mark

Abrasion and scratches Tool marks/dents

REPRODUCIBILITYREPRODUCIBILITYFirst Replica

Fifth Replica

Concern: Repeated Concern: Repeated replication may fill crack replication may fill crack mouthmouthRepeated replicas taken on Repeated replicas taken on several cracked specimens several cracked specimens –– Example: 0.006Example: 0.006--inchinch--long long

surface cracksurface crackNo degradation in crack No degradation in crack detectiondetectionNo debrisNo debris

50 mm

50 mm

APPLICATIONAPPLICATIONReplicaReplica--based based inspection method inspection method approved for use on approved for use on flight hardwareflight hardwareFound 55 cracks in 3 Found 55 cracks in 3 orbiters (684 slots)orbiters (684 slots)–– Ranging from 0.004 Ranging from 0.004

surface crack to 0.040 surface crack to 0.040 inchinch--deep through crackdeep through crack

–– All were missed by eddy All were missed by eddy current inspectioncurrent inspection

Repair confirmationRepair confirmation

OTHER APPLICATIONSOTHER APPLICATIONS

ReplicaReplica--based crack inspection may be based crack inspection may be wellwell--suited for other applicationssuited for other applications–– Improved crack detection could make Improved crack detection could make

damage tolerance life management damage tolerance life management practical for additional componentspractical for additional components

Rotorcraft ?Rotorcraft ?Propellers ?Propellers ?HCF engine components?HCF engine components?

PROS AND CONSPROS AND CONS

Much better Much better resolution than resolution than traditional NDEtraditional NDELittle training Little training required to make required to make replicasreplicasLimited equipment Limited equipment needed in fieldneeded in field

More labor intensive More labor intensive than traditional NDEthan traditional NDELimited to surface Limited to surface flawsflawsDependent on surface Dependent on surface conditionconditionLimited to small areasLimited to small areasNo immediate No immediate feedbackfeedback

PROS CONS

SUMMARYSUMMARY

Analysis of siliconeAnalysis of silicone--based replicasbased replicas–– Find cracks below 0.005 inchesFind cracks below 0.005 inches–– Find pits/defects down to 0.001 inchesFind pits/defects down to 0.001 inchesMethod approved for use on flight Method approved for use on flight hardwarehardware–– Found 55 cracks in 3 orbiters (684 slots)Found 55 cracks in 3 orbiters (684 slots)–– Identified unacceptable levels of damageIdentified unacceptable levels of damage–– Repair confirmed by second round of Repair confirmed by second round of

replicasreplicas

NATIONAL INSTITUTE FOR AVIATION RESEARCH

Wichita State University

30 November 2005 2005 USAF ASIP Conference 1

Destructive Evaluation for Detection of Destructive Evaluation for Detection of Cracks in Lower Wing Stringers and Cracks in Lower Wing Stringers and

Skin of CSkin of C--130E Center Wing Box130E Center Wing Box

USAF Aircraft Structural IntegrityProgram (ASIP) Conference

29 November – 1 December 2005




Introduction• Through the USAF Air Vehicle Health Management program, the Aging Aircraft

Support Squadron in coordination with the 330th Tactical Airlift Support Group selected a decommissioned C-130E aircraft to evaluate its structural integrity.

• The center wing box left and right lower wing skin sections, of the subject C-130E aircraft, were inspected and evaluated by S&K Technologies (SKT), along with the National Institute for Aviation Research (NIAR) and the Center for Aircraft Structural Life Extension (CAStLE).

• Following an ultrasonic inspection of the lower wing skins and stringers, the structure was inspected utilizing conventional nondestructive testing (NDT) such as:

– Visual testing (VT),– Surface scan eddy current (SSEC),– Bolt hole eddy current (BHEC), and– Liquid penetrant inspection (LPI) techniques.

• The destructive evaluation of the left and right lower wing skins and stringers from Wing Station (WS) 68 to WS 206 and from stringer 12 (S-12) to S-24 involved the following tasks:

– Complete surgical disassembly of the wing sections,– Performance of the conventional NDT techniques, and– Failure analysis on selected crack indications.




Inspection Plan• Due to the significance of this critical aircraft section, the complete

disassembly and secondary NDT testing were all scheduled for completion within an eight-week time frame.

• The sequence of disassembly and inspections outlined and followed by NIAR was as follows:

1. Remove all fasteners at locations identified with an indication from the ultrasonic inspection.

2. Perform BHEC inspection on all fastener locations identified in the ultrasonic inspection while the structure is still assembled.

3. Disassemble structure completely and chemically strip all parts.4. Perform LPI on 100% of the structure including all skin and stringer

surfaces. 5. Conduct close VT on 100% of the structure including all holes using 10X

magnification. 6. Conduct SSEC and/or BHEC inspections on all indications identified

from fluorescent PT and/or VT inspections.7. Record all indications.

• Due to unexpected test failures during step 4, which will be discussed later, modifications to the inspection plan were implemented. As a result, over 17,000 bolt holes required BHEC inspection instead of LPI as originally planned.




Initial BHEC Inspections• BHEC INSPECTIONS PRIOR TO

DISASSEMBLY– Performed on selected fastener hole

locations with indications from an automated ultrasonic inspection technique

– Performed on select locations between LWS 130 and LWS 160 by removing the fasteners and inspecting the holes with the rest of the structure intact.

– 72 of 138 of the identified locations had BHEC indications

– Possible reasons for indications:• Cracks• Corrosion in the holes • Sealant

LWS 106

Areas Marked by Ultrasonic InspectionInitial Bolt Hole Eddy Current Indications

LWS 148




Liquid Penetrant Inspections• Once the LPI process began, a noticeable

fluorescent background was present on the surface of the skins, which reduced contrasting and could have masked potential discontinuity indications.

• The chemical stripping process that was used for the LPI preparation was ineffective in removing all organic and inorganic surface coatings. A type of anodize coating remained, covering all inner skin surfaces, and sealant was found around many of the fastener holes.

• As a result, LPI was not an effective testing method so BHEC inspection was conducted on every hole on each skin panel and stringer, requiring inspection of about 17,600 fastener hole locations.




Close Visual Inspections• Close visual inspections were performed with a 10x

microscope on the lower skins and stringers of both the right and left wings in order to identify any visible corrosion or cracks.

• Corrosion was characterized by depth and area of corrosion while any cracks found were identified with their length.

• To facilitate the inspections, the left and right lower wing skins were divided into six sections:– LH WS 68-106 (LH Inboard) − RH WS 68-106 (RH Inboard)– LH WS 106-148 (LH Midboard) − RH WS 106-139 (RH Midboard)– LH WS 148-206 (LH Outboard) − RH WS 139-206 (RH Outboard)




Severe Corrosion in Left Outboard Section

16.41 x 0.10.0320.195S-20 B27Lwr. Surf.

Skin L-panel 2 WS 148-206

88.21.8 x 0.40.1230.1395S-24 B26-B29Upr. Surf.


23.628 x 10.0250.106B2-B29Upr. Surf.

S-23 WS 148-206

15.930 x 10.0170.107A1-A30Upr. Surf.

S-24 WS 148-206

100.030 x 20.1070.107B1-B29Upr. Surf.

S-24 WS 148-206

17.90.35 x 0.350.0250.1395S-23 A26Lwr. Surf.


32.30.6 x 0.60.0450.1395S-23 B26Upr. Surf


76.70.6 x 0.40.1070.1395S-24 B26-B29Lwr. Surf.


Percent Thickness

Lost

Area(inches)

CorrosionDepth (inches)

Part Thickness (inches)

Hole LocationPart Number




Left Upper Skin Surface at Stringer 24 Corrosion Findings

Upper Skin Surface @ S-24LWS 148-206 Holes B26-B29




Left Stringer 24 Upper SurfaceCorrosion Findings

Upper Surface Stringer 24LWS 148-206 Holes B1-B29




Severe Corrosion in Left Mid/Inboard Sections

Percent Thickness

Lost

Area (inches)

Corrosion Depth (inches)


Hole LocationPart Number

17.02.2 x 0.40.0340.2S-24 B30-B32 Upr. Surf.

Skin L-Panel 3 WS 68-104.5

27.12.35 x 0.650.0590.218S-24 B4-B7 Upr. Surf.

Skin L-Panel 3 WS 68-104.5

46.737 x 0.90.0500.107B1-B36 Upr. Surf.

S-24 WS 68-104.5

50.735.5 x 20.0540.1065B1-B36Upr. Surf.

S-24 WS 106-148




Stringer Cracks on RWS 140-207

Stringer 15 Lower Surface RWS 140-207 Hole B41

1.837”

AFT

OTB

Stringer 18 Lower Surface RWS 140-207 Hole A6

0.799”AFT

OTB




Types of Post-Teardown BHEC Indications

Crack Indication from Standard

High Confidence Crack Indications Oriented

Forward and Aft

High Confidence Crack Indication

Lower Confidence Indication with High

Signal to Noise Ratio




Post-Teardown BHEC Inspections

4

6130

40

49

53

0

20

40

60

80

100

120

LWS 68-106 LWS 106-148 LWS 148-206

Final BHEC Skin Final BHEC Stringers

20 9

663638

53

020406080

100120140

RWS 68-106 RWS 106-139 RWS 139-206


• 462 holes with BHEC indications distributed over both the left and right lower wing skins and stringers located between WS68 and WS206.

– 240 indications in the left wing sections– 222 indications in the right wing sections.– 58% of the indications found in stringers.

• Some indications could have been due to the sealant and anodize remaining on the wing surfaces after the paint stripping process.

• Despite these factors, a number of fastener holes were found with high confidence level indications in either the skins or stringers.




Left Wing Inboard Section WS 68-106Fastener #1

S-12S-24Automated Ultrasonic Inspection Indications Initial Bolt Hole Eddy Current Indications Final Bolt Hole Eddy Current Skin Indications Bolt Hole Eddy Current Stringer Indications Indications with High Levels of Confidence in the Findings




Left Wing Midboard Section WS 106-148

Automated Ultrasonic Inspection Indications Initial Bolt Hole Eddy Current Indications Final Bolt Hole Eddy Current Skin Indications Bolt Hole Eddy Current Stringer Indications Indications with High Levels of Confidence in the Findings




Left Wing Outboard Section WS 148-206





Right Wing Inboard Section WS 68-106

S-12S-24 Fastener #1Automated Ultrasonic Inspection Indications Initial Bolt Hole Eddy Current Indications Final Bolt Hole Eddy Current Skin Indications Bolt Hole Eddy Current Stringer Indications Indications with High Levels of Confidence in the Findings




Right Wing Midboard Section WS 106-140





Right Wing Outboard Section WS 140-206





Multiple Indications at Adjacent Locations

• 80 locations in the skins and 111 locations in the stringers were identified with high confidence of a crack indication.

• Numerous areas in several wing sections had multiple indications in the skin or stringer at adjacent hole locations along the same stringer. – In the LWS 106-148 Midboard Section:

• 20 indications in the skin along S-16 from fastener location C-2 to C-37• 16 indications in the skin along S-16 from fastener location D-3 to D-34• 7 indications in the stringer along S-21 from fastener location B-19 to B-32

– In the LWS 148-206 Outboard Section:• 8 indications in the skin along S-16 from fastener location B-34 to B-49• 7 indications in the stringer along S-17 from fastener location B-12 to B-25• 8 indications in the stringer along S-18 from fastener location B-1 to B-14




Failure Analysis• Failure analysis is being performed on fastener hole locations with

indications from both the automated ultrasonic inspection and the BHEC inspection.

– Ultrasonic inspection spanned 30 inches from WS 130 to WS 160– BHEC inspection spanned 138 inches from WS 68 to WS 206.

• Failure analysis results to date showed that a significant number of indications were actually not confirmed as cracks.– 73 of the 93 locations that have been examined showed no cracks.– Failure analysis is still pending on 47 locations.

• The number of locations with no cracks was somewhat anticipated due to several factors:– Sensitivity level at which the BHEC unit was set– Edge burrs– Elongated holes (causing lift-off problems)– Corrosion in some hole– Threading of the holes (due to rivet removal)– Residual sealant remaining on the back surface of the panels




Failure Analysis Findings

9

27

10

6

4

0

9

5

9

0

5

10

15

20

25

30

35

40

LWS 68-106 LWS 106-148 LWS 148-206

No Cracks Cracks Pending

7

15

5

4

6

0

11

11

2

0

5

10

15

20

25

30

35

RWS 68-106 RWS 106-139 RWS 139-206





Summary• The structural integrity of a C-130E center wing box was inspected and evaluated

for cracks and corrosion utilizing conventional NDT techniques.• The destructive evaluation involved the following:

– Complete disassembly of the left and right lower wing skin sections and stringers from WS 68 to WS 206 and from S-12 to S-24,

– Performance of the conventional NDT techniques, and– Failure analysis on selected indications.

• Visual inspections results showed the following:– Extensive corrosion on the left wing sections, mostly at stringer 24, with several

locations having more than 10% thickness loss due to corrosion.– Two visible cracks were found on stringers in the right wing outboard section.

• Due to difficulties in performing LPI on the wing section, over 17,000 fastener holes were inspected in the skins and stringers after disassembly using BHEC.

– 80 locations in the skins and 111 locations in the stringers were identified with high confidence of a crack indication.

– Numerous areas in several wing sections had multiple indications in the skin or stringer at adjacent hole locations along the same stringer.

• Only 20 of the 93 locations examined in the failure analysis had an actual crack. Although a significant number of NDT indications were observed, the actual number of cracks confirmed in the failure analysis was far fewer.




Conclusions• Macroscopic and microscopic analysis of fastener holes with NDT

indications is paramount to “closing the loop” on the actual aircraft / fleet structural integrity evaluations

• Failure analysis is a step forward in establishing a “field based”probability of detection for the NDT techniques employed.

• While the vast majority of NDT indications were not confirmed with the failure analysis, the few cracks found in conjunction with the extensive amount of corrosion present does confirm that cracks and corrosion played a major role in the aging of this aircraft.

• The destructive evaluation provided a comprehensive examination of the critical structure on a high-time aircraft, thereby providing the C-130 System Program Office with valuable data for the assessment of the structural integrity.

• Such information may assist the USAF in determining the airworthiness of their fleet of C-130s.

1

Destructive Evaluation for Detection of Cracks in Lower Wing Stringers and Skin of C-130E Center Wing Box

Dr. Dale Cope and Anthony Alford

Aging Aircraft Laboratory, National Institute for Aviation Research Wichita State University, Wichita, Kansas

Michael Stuerman S&K Technologies

Dayton, Ohio

ABSTRACT

Non-destructive testing (NDT) techniques are an essential part of a structural integrity program. As part of the U. S. Air Force Air Vehicle Health Management program, the center wing box of a decommissioned C130E aircraft was selected for a teardown and failure analysis in order to determine the extent of structural damage in this aircraft section. Following an ultrasonic inspection of lower wing skins and stringers, S&K Technologies, along with the National Institute for Aviation Research and the Center for Aircraft Structural Life Extension, collectively identified the extent of damage in the lower wing skin and stringers of the center wing box through a destructive evaluation. This evaluation involved the bolt-hole eddy current (BHEC) inspection of select fastener locations identified in the ultrasonic inspection, complete disassembly of the lower skins and stringers, visual and liquid penetrant inspection of the structure, and BHEC inspection of all fasteners holes in both the skins and stringers. These inspections were followed up by a failure analysis of crack indications in order to definitively determine the extent of damage in the structure. The results from the visual inspections showed extensive corrosion on the left wing sections, mostly at stringer 24, with several locations having more than 10% thickness loss due to corrosion. Two visible cracks were discovered in separate stringers on the right wing outboard section. Due to difficulties in performing LPI on the wing section, over 17,000 fastener holes were inspected in the skins and stringers after disassembly using BHEC, and 191 locations were identified with high confidence of a crack indication. Numerous areas had multiple indications in the skin or stringer at adjacent hole locations along the same stringer. Of the 93 locations examined in the failure analysis, only 20 were found to have an actual crack. These results illustrate that macroscopic and microscopic analysis of fastener holes with NDT indications is paramount to “closing the loop” on actual aircraft / fleet structural integrity evaluations and a step forward in establishing a “field based” probability of detection for the NDT techniques employed. While the vast majority of NDT indications on the C-130E center wing box were not confirmed with the failure analysis, the few cracks found in conjunction with the extensive amount of corrosion present did confirm that cracks and corrosion played a major role in the aging of this aircraft. The destructive evaluation provided a comprehensive examination of the critical structure on a high-time aircraft, thereby providing the C-130 System Program Office with valuable data for the assessment of the structural integrity. Keywords: stress, fracture, conventional nondestructive testing, visual testing (VT), surface scan eddy current (SSEC), bolt hole eddy current (BHEC), liquid penetrant inspection (LPI), failure analysis

2

INTRODUCTION As current commercial and military aircrafts begin to reach their design service goal (or objective), service failures due to stress, overload, and fatigue are probable. All aircraft are subject to stress and fatigue that can cause permanent structural damage to an aircraft. The ability to detect flaws in primary structural members before the defect reaches a critical size is an essential part of ensuring the extended life of an aircraft. The United States Air Force (USAF) has always paid particular attention to safety and longevity of their aircraft. Through the USAF Air Vehicle Health Management program, the Aging Aircraft Support Squadron in coordination with the 330th Tactical Airlift Support Group selected a decommissioned C-130E aircraft to evaluate its structural integrity. As part of this evaluation, the center wing box left and right lower wing skin sections, of the subject C-130E aircraft, were inspected and evaluated by S&K Technologies (SKT), along with the National Institute for Aviation Research (NIAR) and the Center for Aircraft Structural Life Extension (CAStLE). The structure was inspected utilizing conventional nondestructive testing (NDT) such as visual testing (VT), surface scan eddy current (SSEC), bolt hole eddy current (BHEC), and liquid penetrant inspection (LPI) techniques. The destructive evaluation involved the complete surgical disassembly of the wing sections, performance of the conventional NDT techniques, and failure analysis on the left and right lower wing skins and stringers from Wing Station (WS) 68 to WS 206 and from stringer 12 (S-12) to S-24. NIAR performed all the conventional NDT testing; SKT and CAStLE conducted the failure analysis through fractographic examinations on identified crack indications from the conventional NDT testing; and SKT correlated the results between the NDT techniques.

INSPECTION PLAN Due to the significance of this critical aircraft section, the complete disassembly and secondary NDT testing were all scheduled for completion within an eight-week time frame. The sequence of disassembly and inspections outlined and followed by NIAR was as follows:

1. Remove all fasteners at locations identified with an indication from the ultrasonic inspection.

2. Perform BHEC inspection on all fastener locations identified in the ultrasonic inspection while the structure is still assembled.

3. Disassemble structure completely. 4. Chemically strip all parts. 5. Perform LPI on 100% of the structure including all skin and stringer surfaces. 6. Conduct close VT on 100% of the structure including all holes using 10X magnification. 7. Conduct SSEC and/or BHEC inspections on all indications identified from fluorescent

PT and/or VT inspections. 8. Record all indications.

Due to unexpected test failures during step 5, which will be discussed later, modifications to the inspection plan were implemented. As a result, over 17,000 bolt holes required BHEC inspection instead of LPI as originally planned. Non-destructive inspections were performed on the six lower wing skin sections and their respective stringers. All inspections were performed by certified NDT Level II technicians and qualified visual inspectors. All bolt hole eddy current inspections were performed according to an approved USAF procedure utilizing an automated scanning device to inspect fastener holes.

3

CONVENTIONAL NONDESTRUCTIVE INSPECTIONS

BHEC INSPECTIONS PRIOR TO DISASSEMBLY Prior to disassembly, BHEC inspections were performed on selected fastener hole locations, which were indications identified by an automated ultrasonic inspection technique performed only between Wing Station (WS) 130 and WS 160 on the left and right lower wing surfaces. These inspections were performed by removing only those select fasteners and inspecting the holes with the rest of the structure intact. Then, the skins and stringers of each wing section were fully disassembled and paint stripped in order to perform the next steps.

LIQUID PENETRANT INSPECTIONS Fluorescent LPI was performed on the left wing mid-board and outboard sections using an approved USAF procedure. This procedure was based on a method D (hydrophilic removable) penetrant inspection process. The process required a minimum dwell time of one hour, unless stress corrosion cracking was suspected, in which case a four hour dwell time was used. The LPI process depends upon the ability of the penetrant to seep into a discontinuity. Any factor that interferes with the entry or exit of the penetrant reduces the effectiveness of the inspection. Once the LPI process began, a noticeable fluorescent background was present on the surface of the skins, which reduced contrasting and could have masked potential discontinuity indications. In order for LPI to be a high-quality inspection, the part’s surface must be clean and free of organic or inorganic contaminants. The chemical stripping process that was used for the LPI preparation was ineffective in removing all organic and inorganic surface coatings. A type of anodize coating remained, covering all inner skin surfaces, and sealant was found around many of the fastener holes. Figure 1 shows an example of this problem. As a result, LPI was not an effective testing method so BHEC inspection was conducted on every hole on each skin panel and stringer. This modification to the inspection plan created a significant increase in the number of holes that required BHEC inspection. Of the approximately 17,600 fastener hole locations in the skins and stringers, all of these locations were inspected using the BHEC inspection method.

Figure 1. Fluorescent Penetrant Background on Left Wing Skin Panel WS 148-206

CLOSE VISUAL INSPECTIONS Close visual inspections were performed with a 10x microscope on the lower skins and stringers of both the right and left wings in order to identify any visible corrosion or cracks. Corrosion was characterized by depth and area of corrosion while any cracks found were identified with their length.

4

To facilitate the inspections, the left and right lower wing skins were divided into six sections:

• LH WS 68-106 (LH Inboard) • RH WS 68-106 (RH Inboard) • LH WS 106-148 (LH Midboard) • RH WS 106-139 (RH Midboard) • LH WS 148-206 (LH Outboard) • RH WS 139-206 (RH Outboard)

Since inspection results are primarily tracked by fastener location, all fastener locations for the six sections were identified according to which row they were in on their respective stringer, labeled alphabetically starting with the forward side of the stringer. Then, fastener locations were given a number to distinguish their location on the skin or stringer from inboard to outboard. The fastener numbering for each lower wing section began with number 1 so the wing section and alpha-numeric designation together uniquely identified fastener locations as to where inspection indications are located. Extensive corrosion was discovered on the left wing sections, mostly at stringer 24, with several locations having more than 10% thickness loss due to corrosion, as listed in Table 1. Only light corrosion, with thickness loss of less than 2%, was noted on the right wing sections.

Table 1. Corrosion Identified on Left Wing Skin Panels And Stringers

Skin/Stringer Wing Section Hole Location


Corrosion Depth

(inches)

Area (square inches)

% Thickness

Loss Skin Panel

LWS 148-206 S-24 B26-B29

Upr. Surf. 0.1395 0.123 0.72 88.2

Skin Panel LWS 148-206

S-24 B26-B29 Lwr. Surf. 0.1395 0.107 0.24 76.7


S-23 B26 Upr. Surf 0.1395 0.045 0.36 32.3


S-23 A26 Lwr. Surf. 0.1395 0.025 0.1225 17.9


S-20 B27 Lwr. Surf. 0.195 0.032 0.1 16.4

S-24 LWS 148-206

B1-B29 Upr. Surf. 0.107 0.107 60.0 100.0

S-24 LWS 148-206

A1-A30 Upr. Surf. 0.107 0.017 30.0 15.9

S-23 LWS 148-206

B2-B29 Upr. Surf. 0.106 0.025 28.0 23.6

S-24 LWS 106-148

B1-B36 Upr. Surf. 0.1065 0.054 71.0 50.7

S-24 LWS 68-106

B1-B36 Upr. Surf. 0.107 0.050 33.3 46.7


S-24 B4-B7 Upr. Surf. 0.218 0.059 1.5275 27.1


S-24 B30-B32 Upr. Surf. 0.2 0.034 0.88 17.0

5

Some of the most severe areas of corrosion on the left wing sections involved corrosion depths of 46.7% to 100% through the thickness of the skin or stringer. As shown in Figures 2 and 3, two areas of corrosion on the left wing outboard section at S-24 had thickness loss of 88.2% and 76.7%, respectively, on the upper and lower skin surfaces. These two areas of corrosion were in the same region, however they did not overlap each other, and at no location did the corrosion go through the entire skin. Two other areas of severe corrosion were found on the left outboard skin panel at specific fastener locations on S-23 and S-20.

Figure 2. Corrosion at Upper Skin Surface, LWS 148-206 S-24

Figure 3. Corrosion at Lower Skin Surface, LWS 148-206 S-24

6

Figure 4 shows the significant corrosion found on the upper surface of S-24, common to the outboard skin. This corrosion encompassed an area of 60 square inches. The maximum depth corresponded to a 100% thickness loss at one small location with other areas severely corroded as well. Another area of corrosion on the outboard section of S-24 encompassed an area of 30 square inches and had a corrosion depth through 15.9% of the thickness. The outboard section of S-23 also had an area of corrosion encompassing an area of 28.0 square inches and having a corrosion depth through 23.6% of the thickness.

Figure 4. Corrosion on LWS 148-206 Stringer S-24 Upper Surface

Extensive corrosion was also found on the upper surface of S-24 on the mid-board (WS 106-148) and inboard (WS 68-106) sections, similar to the corrosion found on the outboard section of the same stringer. On the mid-board section of S-24, the corrosion encompassed an area of 71 square inches, and the maximum depth of the corrosion was 50.7% through the thickness. Again, the upper surface of S-24 was found to be heavily corroded, as shown in figure 5. This widespread corrosion was located around several holes, covered an area of 33.3 square inches, and had a maximum depth that equated to a 46.7% material loss. The inboard skin panel also had two localized areas of severe corrosion, as listed at the end of table 1.

Figure 5. Corrosion on LWS 68-106 Stringer S-24 Upper Surface

7

During the close visual inspections, no cracks were noted on the left wing sections, and only two visible cracks were discovered on the right wing sections. Figure 6 shows a crack found on stringer 15 on the outboard section. This crack measured 1.837 inches in total length, and it was through the thickness of the material, which was 0.103 inches thick. Another crack measuring 0.799 inches was found on stringer 18 on the outboard section, as shown in Figure 7. It was also through the thickness of the material, which was 0.108 inches thick.

Figure 6. Crack (1.837”) on Stringer 15 Lower Surface RWS 139-206

Figure 7. Crack on Stringer 18 Upper Surface RWS 139-206, Hole A6-A7

0.799”

A6

A7

FWD

OTB

AFT

OTB

8

POST-TEARDOWN BHEC INSPECTIONS Following the complete disassembly of all six wing skin panels, BHEC was performed on all fastener locations in both the wing skin and stringers. Because of the large number of indications and based on guidance from the C-130 System Program Office, inspectors proceeded to group indications as either high or low level confidence indications based on BHEC signal amplitude and signature. This grouping would serve to prioritize fastener holes for follow-on failure analysis. Examples of the BHEC screen displays illustrating this grouping criteria can be viewed in figures 8 through 11. Figure 8 shows the eddy current signal obtained from the 0.030 inch crack standard while figure 9 shows the eddy current signal of a crack found by visual inspection. This crack was noted on both sides of a hole on a stringer. The eddy current signal for a high confidence indication is shown in figure 10. This indication closely resembles both the crack standard and actual crack signals. A high confidence indication should have a high signal to noise ratio as seen in figures 8 through 10. A low confidence indication is shown in figure 11, illustrating a low signal to noise ratio, which could indicate corrosion or damage in the hole. The low confidence indication is not typical, however, of an actual crack. All indications on the skins and stringers were classified as either high or low confidence indications by the qualified NDI Level II inspector.

Figure 8. Eddy Current Signal on the 0.030-in Crack Standard

Figure 9. Eddy Current Signal of a Crack Found with Visual Inspection

9

Figure 10. Eddy Current Signal for a Highly Confident Indication

Figure 11. Eddy Current Signal for a Low Confidence Indication

The BHEC inspection yielded 462 holes with indications distributed over both the left and right lower wing skins and stringers located between WS68 and WS206. Figures 12 and 13 show the number of BHEC indications found in the skins and stringers of each wing section, with a total of 240 BHEC indications in the left wing sections and 222 BHEC indications in the right wing sections. The indications were located throughout the six wing sections with 58% of the indications being found in stringers. Inspectors suspected that a good portion of the indications could be due to the sealant and anodize remaining on the wing surfaces after the paint stripping process. Despite these factors, a number of fastener holes were found with high confidence level indications in either the skins or stringers.

10

4

6130

40

49

53

0

20

40

60

80

100

120

LWS 68-106 LWS 106-148 LWS 148-206


Figure 12. Number of BHEC Indications in Left Wing Sections, WS 68-206

20 9

663638

53

020406080

100120140

RWS 68-106 RWS 106-139 RWS 139-206


Figure 13. Number of BHEC Indications in Right Wing Sections, WS 68-206

RESULTS OF NON-DESTRUCTIVE INSPECTIONS In order to better understand the significance of all the NDI results, pictorial overlays were generated for each wing section showing the locations of the various indications. Figures 14-19 show the indications from the automated ultrasonic inspection, initial BHEC, post-teardown BHEC on the skins, and BHEC on the stringers. High confidence indications, as defined previously, are also indicated on the pictorial overlays. In each of the overlays, the following symbols were used to mark the location of the various indications.


11

LEFT WING INSPECTION RESULTS Figure 14 compares the indications on the left wing inboard section, while figures 15 and 16 show the left wing mid-board and outboard sections, respectively.

Figure 14. Indications on Left Wing Inboard Section, WS 68-106 (Panel 1)

Figure 15. Indications on Left Wing Midboard Section, WS 106-148 (Panel 2)

S-24 S-12

S-12S-24

12

Figure 16. Indications on Left Wing Outboard Section, WS 148-206 (Panel3)

RIGHT WING INSPECTION RESULTS Similar to the left wing, figure 17 provide a comparison of NDT indications on the right wing inboard section. Likewise, figure 18 detail the indications on the right wing mid-board section and figure 19 for the outboard section.

Figure 17. Indications on Right Wing Inboard Section, WS 68-106 (Panel 1)

S-12 S-24

S-24 S-12

13

Figure 18. Indications On Right Wing Midboard Section, WS 106-139 (Panel 2)

Figure 19. Indications on Right Wing Outboard Section, WS 139-206 (Panel 3)

S-12 S-24

S-12 S-24

14

SUMMARY OF BHEC INSPECTION RESULTS Of the approximately 17,000 fastener hole locations inspected in the skins and stringers using BHEC, 80 locations in the skins and 111 locations in the stringers were identified with high confidence of a crack indication. Numerous areas in several wing sections had multiple indications in the skin or stringer at adjacent hole locations along the same stringer. Some of the areas with 7 or more adjacent indications included the following:

• In the LWS 106-148 Midboard Section: o 20 indications in the skin along S-16 from fastener location C-2 to C-37 o 16 indications in the skin along S-16 from fastener location D-3 to D-34 o 7 indications in the stringer along S-21 from fastener location B-19 to B-32

• In the LWS 148-206 Outboard Section: o 8 indications in the skin along S-16 from fastener location B-34 to B-49 o 7 indications in the stringer along S-17 from fastener location B-12 to B-25 o 8 indications in the stringer along S-18 from fastener location B-1 to B-14

• In the RWS 68-106 Inboard Section: o 9 indications in the skin along S-14 from fastener location B-10 to B-37

• In the RWS 106-139 Midboard Section: o 7 indications in the stringer along S-16 from fastener location D-5 to D-23

• In the RWS 139-206 Outboard Section o 8 indications in the skin along S-14 from fastener location A-6 to A-31

FAILURE ANALYSIS RESULTS Failure analysis is being performed on fastener hole locations with indications from both the automated ultrasonic inspection and the BHEC inspection. The ultrasonic inspection spanned 30 inches from WS 130 to WS 160, and the area inspected with the BHEC inspection spanned across approximately 138 inches, which represented WS 68-206. As represented in Figures 20 and 21, the failure analysis results to date showed that a significant number of indications were actually not confirmed as cracks, with 73 of the 93 locations examined showing no cracks. Failure analysis is still pending on 47 locations. The number of locations with no cracks was somewhat anticipated due to several factors including the sensitivity level at which the BHEC unit was set, edge burrs, elongated holes (causing lift-off problems), corrosion in some hole, threading of the holes (due to rivet removal), and of course, residual sealant remaining on the back surface of the panels. All of these imperfections can cause “crack-like” indications. Some false positives indication may have been eliminated, if the holes would have been reamed prior to inspection; however, 400-600 grit emery cloth paper was the only permissible pre-inspection hole “cleaning” treatment as established by the SKT and the USAF. This restriction was established in order to prevent the loss of damage to confirmed crack features such as nucleation sites or fracture surfaces.

15

9

27

10

6

4

0

9

5

9

0

5

10

15

20

25

30

35

40

LWS 68-106 LWS 106-148 LWS 148-206


Figure 20. Failure Analysis Finding on Left Lower Wing Skin and Stringers

7

15

5

4

6

0

11

11

2

0

5

10

15

20

25

30

35

RWS 68-106 RWS 106-139 RWS 139-206


Figure 21. Failure Analysis Finding on Right Lower Wing Skin and Stringers

16

SUMMARY AND CONCLUSION Through the USAF Air Vehicle Health Management program, the structural integrity of a center wing box from a decommissioned C-130E aircraft was evaluated. The center wing box is a critical fatigue component for the C-130 since it is structurally more susceptible to the stresses of mission profile and payload. As part of this evaluation, the center wing box left and right lower wing skin sections and stringers, of the subject C-130E aircraft, were inspected and evaluated for cracks and corrosion utilizing conventional NDT techniques. The destructive evaluation involved the complete disassembly of the wing sections, performance of the conventional NDT techniques, and failure analysis on selected indications in the left and right lower wing skins and stringers from WS 68 to WS 206 and from S-12 to S-24. The results from the visual inspections showed extensive corrosion on the left wing sections, mostly at stringer 24, with several locations having more than 10% thickness loss due to corrosion. Two visible cracks were discovered in separate stringers on the right wing outboard section. Each crack was through the thickness of the stringer and longer than a half inch. Due to difficulties in performing LPI on the wing section, over 17,000 fastener holes were inspected in the skins and stringers after disassembly using BHEC. Because of the large number of indications, confidence levels were established based on BHEC signal amplitude and signature. 80 locations in the skins and 111 locations in the stringers were identified with high confidence of a crack indication. Numerous areas in several wing sections had multiple indications in the skin or stringer at adjacent hole locations along the same stringer. Only 20 of the 93 locations examined in the failure analysis had an actual crack. Although a significant number of NDT indications were observed, the actual number of cracks confirmed in the failure analysis was far fewer. This “false call rate” can be explained by the guidelines established with regard to hole clean-up and by the degree of observed mechanical damage in the fastener holes. Macroscopic and microscopic analysis of fastener holes with NDT indications is paramount to “closing the loop” on the actual aircraft / fleet structural integrity evaluations and a step forward in establishing a “field based” probability of detection for the NDT techniques employed. While the vast majority of NDT indications were not confirmed with the failure analysis, the few cracks found in conjunction with the extensive amount of corrosion present does confirm that cracks and corrosion played a major role in the aging of this aircraft. The destructive evaluation provided a comprehensive examination of the critical structure on a high-time aircraft, thereby providing the C-130 System Program Office with valuable data for the assessment of the structural integrity. Such information may assist the USAF in determining the airworthiness of their fleet of C-130s.

1

ASSESSMENT OF THE CAPABILITIES AND READINESS OF NDI METHODS FOR

SUBSURFACE CRACK DETECTION IN BUILT-UP STRUCTURE

David PiotrowskiSenior Engineer - Enabling Technologies

Delta Air Lines

Presented at the 2005 ASIP Conference

Memphis, TNNovember 30, 2005

2

AcknowledgementsThis project was conducted under FAA R&D Contract No. DTFA03-02-C-00044, “DESTRUCTIVE EVALUATION AND EXTENDED FATIGUE TESTING OF A RETIREDPASSENGER AIRCRAFT (B727)”.

FAA Technical CenterJohn Bakuckas, David Galella, Paul Swindell, Amlan Duttchoudhury, Doug Koriakian

Delta Air LinesJohn Bohler, Richard Watkins, Aubrey Carter, David Steadman, Ramesh Ramakrishnan, Doug Jury, Mark Boudreau

Drexel UniversityBao Mosinyi, Jonathan Awerbuch, Alan Lau, T.M. Tan

FAA AANC (Sandia)Mike Bode, David Moore, Floyd Spencer

Gov’t, Academia, Industry partnership

3

Outline of Presentation

• Introduction & Project Overview

• Field, Pre-teardown, & Emerging Inspections

• Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) NDT Support

• Assessment Protocol, Methodology

• NDT Rating System

• Explanation of ‘sites’

• Probability Of Detection Results

• NDT Report Card

• Summary and Future plans

F1F1

FT1FT1(F3 (F3 OppOpp))

FT2FT2F4 (Opp)F4 (Opp) FT3FT3

F5 (Opp)F5 (Opp)

F2aF2aF2bF2b

F6F6FT4 (Opp)FT4 (Opp)

4

Overview and Background• This program is used to establish procedures and guidelines for teardown

activities for use in precluding Widespread Fatigue Damage (WFD).– Assess the state of damage at the Design Service Goal (DSG)– Lack of information in the public domain (many military, OEM programs)– Fine-tune analytical tools used in engineering analysis

• The program is focused on the fuselage of a retired Delta 727-200 aircraft. The objectives are:

– to assess the capabilities of existing and emerging NDT methods to detect multiple site damage (MSD);

– to characterize the state of MSD in fuselage structure of a retired transport airplane at Design Service Goal;

– to advance the state of MSD in selected sections through extended fatigue testing;– to develop analysis methods that can correlate the state of MSD at any point in time.

• The output for the project will help several areas in the industry.– Guidelines to develop, assess and approve programs to preclude Widespread Fatigue

Damage (WFD), supporting WFD rulemaking activities– Calibration and validation of MSD assessment methods– Guidelines for conducting a teardown– Guidelines for Stress Spectra development– Evaluation of field capability of conventional and emerging NDT technologies

Significant benefit to Industry & Regulators

5

Program Overview• Background and Preparation

– Selection of aircraft/areas– Aircraft information report– Specimen removal & test plan

• Inspection Capability Assessment– Conduct field inspection– Conduct post-removal inspection– Emerging NDT– Assess inspection capability

• Damage Characterization– Crack locations/shapes/orientations– Contributing factors to cracking– Reconstructed crack histories (SEM)

• Data Analysis– Spectra development– Initial flaw distributions

• Extended Fatigue Testing– Preparation of panels– Develop test plans– Conduct fatigue tests

• Documentation and Database

Hoop Load Assembly

Shear Load Fixture

Counter Balance Basket

Counter Balance Pole

Longitudinal Load Assembly

Pressure Box

Shear Water Actuator

0.05"0.05"

Slot surfaceSlot surface

Crack surfaceCrack surfaceHole surfaceCrack frontCrack front

Faying surfaceFaying surface

NDT = only a portion of project scope

6

Conduct Field Inspection• Field inspections simulate actual in-service

maintenance • Detailed visual inspection with photographs to

document and catalog condition of panels:– Cracks, dents, corrosion, preexisting repairs

• NDT for cracks, corrosion and disbonds per standard OEM recommended procedures:

– Airworthiness Directives– Service Bulletin– Boeing NDT Manual

• Field inspections performed to allow comparison to Post-removal/Pre-teardown inspections

Inspections simulate actual in-service maintenance

7

Post Removal (Pre-teardown) Inspection

• Field inspections repeated– Allows for comparison (Human Factors)

• Post-removal visual inspection– Conducted under climate-controlled “shop” conditions

• NDT for cracks, corrosion and disbonds– Conventional procedures

• Emerging methods conducted– Data collected for future analysis (Screen capture)

Inspections repeated under climate-controlled environment

8

Initial Findings using Conventional NDT

Stations MFEC LFEC DVI MFEC LFEC DVI 420-440 1 0 0 1 0 0440-460 1 0 0 1 0 0480-500 1 1 0 0 1 0500-520 3 0 1 7 0 1520-540 8 4 6 8 4 6540-560 12 1 10 11 2 10560-580 6 0 1 12 0 1580-600 10 0 0 13 0 0600-620 5 2 3 4 2 0620-640 5 0 5 6 0 5640-660 0 0 0 2 0 0660-680 4 1 0 2 0 0680-700 1 1 0 1 0 0700-720 6 0 0 8 0 0

720-720A 9 2 3 9 3 3720A-720B 8 1 0 8 1 0720B-720C 13 4 0 14 5 0720C-720D 4 1 0 5 1 0720D-720E 3 0 1 3 0 2720E - 720F 0 0 0 1 0 0

Totals 100 18 30 116 19 28

S- 4R

S –4L

FS 259 FS 720

• Number of fasteners with crack indications using Medium Frequency Eddy Current (MFEC), Low Frequency Eddy Current (LFEC) and Detailed Visual Inspection (DVI)

Field Inspections Post Removal Inspections

Conclusion: Field and Post-removal results are the same;

Candidate aircraft was a good selection

9

Next Steps: Extended Fatigue Testing• The objective is to continue

the growth of Multiple Site Damage in a realistic way.

• Full-Scale AircraftStructural TestEvaluation and Research(FASTER) fixture appliesinternal pressure, plushoop, longitudinal, andshear loads to curvedpanel test articleextracted from crownarea.

Test will provide:• Crack initiation and growth

rates, link-up.• Critical MSD distribution

when FAR's not met.• Cycles from conservative

analytical failure criteria to actual failure.

FASTER fixture prior to panel installation. Shows top of pressure box, hoop loader, and shear fixture counterweight frames

FT2 test panel installed. Remote external cameras shown. Remote internal underwater camera is also in use

Panel installed with shear fixture Exploded view

Unique fixture used to simulate additional service cycles

10

FASTER Test Support

NDT supports FASTER test

• NDT support of FASTER testing includes both “Fielded” and “Emerging” techniques– Internal MFEC – Giant Magnetoresistive Sensor (GMR)– Detailed Visual Inspection – Turbo-Magneto Optical Imaging (T-MOI)– Remote Controlled Cameras – High Frequency Linear Array Ultrasonics– LFEC Sliding Probe – Meandering Winding Magnetometer (MWM)– Rivet Check – Magneto Optical Imaging (MOI)

• Inspections support the different phases of FASTER test– Phase I: Initial crack detection (MSD definition)

• Visual inspections every 500 cycles, NDT interval 2000 cycles • Delta support (site visit) periodically

– Phase II: Crack growth measurements (and monitor new cracks)• Visual/NDT more frequent, NDT every 1000 cycles• Extensive Delta support (site visit)

• All data put into project database(See Steadman Presentation at 10:30)

11

• Characterization of cracks along 4R:– Multiple cracks forming a starburst– Multiple crack origins: rivet hole and

faying surface– Eventually form a contiguous crack– Crack tunneling under clad layer

• Effect on NDT:– Internal Visual/HFEC methods would not

detect until breakthrough (~0.250”)– Consistency of sealant between layers– Orientation affects LFEC sliding probe– Faying surface origin = No BHEC detection

Damage Characterization

Findings significantly affect NDT Inspections

12

Microscopic examination of cracking during Damage Characterization.

Lower row S-4R cracks as depicted by an emerging NDT method - external LFEC on SAIC automated scanner.

External MOI NDT under controlled conditions after panel removal.

Inspection Capability Assessment

• Results of the Field and Pre-teardown inspections to be compared to crack measurements from Damage Characterization.

• Developed system to rate emerging NDT’s readiness for airline use.

Lower row S-4R cracks as depicted by an emerging NDT method - external MOI.

The objective is to assess the capabilities the selected NDT used in this study to find and characterize damage.

Combining Damage Characterization and NDT results

13

Inspection Capability Assessments

Technology Vendor/Process Detailed Visual Generally accepted practices Low Frequency Eddy Current (LFEC) Sliding Probe Boeing NDTM Medium Frequency Eddy Current (MFEC) Boeing NDTM Automatic Couplant Ejection System (ACES) SAIC C-Scan Eddy Current with Sliding Probe Delta Air Lines Conventional Film Radiography Generally accepted practices Digital Radiography Virtual Media Imaging (VMI) Eddy Current Array Probe R/D Tech Eddyscan Nortec (Staveley) Giant Magnetoresistive Sensor (GMR) NASA Langley High Frequency Ultrasonic Array USUT Labs Magneto Optical Imaging (MOI) PRI MAUS Trescan NDT Solutions & Boeing, St. Louis MAUS Rotoscan Boeing, St. Louis Mobile Automated Scanner (MAUS) Rasterscan Boeing, St. Louis Meandering Winding Magneometer (MWM) sensor Jentek Remote Field Eddy Current IMTT Self-Nulling Rotating Probe (Rivet Check) Foerster & FAA AANC Structural Anomaly Mapping (SAM) Honeywell Turbo Magneto Optical Imaging (MOI) PRI & Boeing, Seattle

• 20 techniques evaluated; 2 techniques declined participation• Collaboration of Delta, FAA-AANC, FAA-TC, Boeing selected methods

Variety of NDT Techniques explored, compared to current NDT

14

Protocol for Inspection Assessments• Comparison of techniques made based on Emerging, Pre-teardown (shop) and Field

inspections, FASTER inspections and actual in-service experience.

• Inspection protocol used to establish “level playing field”

– According to the FAA’s Aging Aircraft Nondestructive Validation Center (AANC), the following variables are important to control:

• Location/facility – Lighting, background noise, temperature, comfort

• Inspector – Human factor variables, training and qualification, certification

• Equipment – Variables with calibration, probe types, instrument settings

• Inspection Process Variables – Decision thresholds, scan rate, etc.

– Variables were controlled by:• Location – All Pre-teardown (shop) and Emerging NDT took place in same climate controlled area

• Inspector – Field, Pre-teardown (shop), some FASTER inspections performed by Level IIIs; most Emerging NDT performed by company reps or inventors

• Equipment – Identical calibration procedures and reference standard assures level playing field

• Inspection Process Variables – Each Emerging NDT inspector provided time limit, but allowed tomaintain own pace; decision threshold identical; participants prohibited from validating results

• Emerging NDT focus is on the techniques, not the inspectors– “Capability” versus “Reliability”

Inspection protocol allows for valid comparisons

15

Methodology for Inspection Assessments• Four categories for comparisons:

– Sensitivity: Based on the Probability of Detection for each technique. Values based on “truth” data obtained through Damage Characterization; Awarded from 1-20 points, based on order from worst to best

– Ease-of-use: Calibration and use of the software required for an actual NDT inspector to interpret the results. Observations from Delta Level III personnel & Delta inspectors. Rated on a scale from one to 10, with 10 rated “substantially above current inspections”, five rated “same as current inspections”, and one rated “substantially below current inspections”.

– Speed of the inspection: Initial scanning rate and final data analysis, if separate from the inspection. Rated on same one to 10 scale as above.

– “Fieldability”: Portability of the inspection, & the projected durability to operate in the airline hangar environment (i.e., weight, wires, connections, “drop-ability”, etc.). Rated on same one to 10 scale as above.

• Assessment made with thought toward using Emerging NDT techniques to solve or improve large-area in-service issues: MD88 crown inspections (SSID); B737 mill-line cracking; Scribe lines; Lap joint inspection on other aircraft.

4 categories allow for end-user to weight “preferences”

Based on observations/ experience of Level IIIs &actual inspectors

16

“Single opportunity” =Largest crack at fastener

“Two opportunities” =Largest crack on each side

(fwd/aft) of fastener

Multiple Opportunities =All cracks counted

Factors Affecting Probability of Detection Analysis

!

!

!

!

!

!

• Inspection Sites or “Opportunities”Actual Damage Characterization Results NDT Analysis

Definition of inspection sites affects POD analysis

17

• Flaw size distribution not adequate for some 2-parameter hit/miss POD analysis (default)– Requires always detected large flaws

– Requires always missed small flaws (i.e., data set is not MIL-HDBK-1823 compliant)

– Used 4-parameter, Variable angle or Multi-flaw model fits• FAA-AANC leading effort to determine method which addresses the amount of damage at an inspection site• Ability to differentiate individual flaws leads to algorithms that generate distinct detectable flaws for a given technology

• Also performed some signal analysis (a-hat versus a)

Other factors• ‘Single opportunity’, ‘two opportunity’ or ‘multiple

opportunity’, or combination

• ‘Rejectable’ and ‘non-rejectable’ indications

• Areas scanned (not all techniques scanned same areas)

– Damage Characterization and Panel modification timing

0

10

20

30

40

50

60

70

80

90

100

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00 1.10 1.20 1.30 1.40 1.50

ACTUAL CRACK LENGTH - (Inch)PR

OB

AB

ILIT

Y O

F D

ETEC

TIO

N (%

)

Typical POD Curve with

confidence curve

90/95 curve

90/50 curve

Data set provided some issues for standard 2-parameter hit/miss analysis

Factors Affecting Probability of Detection Analysis

18

Signal Analysis: EDM Notch vs Cracks

EDM Notch vs Cracks: Long-running debate within NDI Community

• EDM Notch often used as Calibration

• But EDM notches are not cracks!

• 3-point calibration (Rummel) discussions

– Instrument linearity– Linear relationship assumed for crack

response (i.e., same crack length as EDM)

• This project had 3-EDM notches in the Calibration Standard provided to all methods

– 0.050”, 0.100”, and 0.150” thru thickness(0.040” skin) 2nd layer EDM Notches

• Signal data collected and archived throughthe database (Contract deliverable)

EDM Notch vs Cracks MFEC Comparison

0

25

50

75

100

125

150

0 0.05 0.1 0.15 0.2 0.25 0.3

Crack Length (in)

Sign

al

EDM Notches

Reject Threshold

Cracks

0.1” EDM Notch equals a 0.140” crack via MFEC

19POD of NDT Techniques

NDT Analysis – POD Results

0.1020.00.129100.082.00.0910.071MFEC - Pre-teardown (Single)

0.00.129100.072.00.0870.074MFEC - Pre-teardown (Double)

0.0870.00.20296.468.00.1030.081MFEC - Field (Single)

0.00.20294.964.00.1130.092MFEC - Field (Double)

0.2130.00.20266.725.0N/A0.428MAUS Rotoscan (Single)

0.3250.00.23746.222.00.7230.357MAUS Rasterscan (Single)

0.2500.00.23739.313.10.4160.286LFEC - Pre-teardown (Single)

0.3000.00.23739.313.10.4170.287LFEC - Field (Single)

0.00.23742.913.00.2470.193Giant Magnetoresistive Sensor (Single)

0.0N/AN/A0.0N/AN/AEddyscan

0.5003 (2.4%)0.26339.322.0N/A0.998DVI - Pre-teardown (Single)

0.3003 (1.4%)0.26338.524.2N/A1.042DVI - Pre-teardown (Double)

0.3503 (2.4%)0.26332.118.2N/A1.314DVI - Field (Single)

0.2173 (1.4%)0.26330.819.1N/A0.934DVI - Field (Double)

0.1800.00.20466.757.20.3280.191Digital Radiography (Single)

0.00.20470.862.60.3400.207Digital Radiography (Double)

3 (2.4%)0.26353.626.00.5130.259C-scan Eddy Current (Single)

0.00.2630.00.0N/AN/AConventional Film X-ray

0.2400.00.23717.916.0N/AN/AArray Eddy Current (Single)

0.2655 (4.2%)0.24932.115.0N/A1.083ACES (Single)

flaw size(rate)Missed(>0.150)(All flaws)flaw sizeflaw size

detectedfalse callsFlawdetecteddetected90/950.9

AANC 4-parameter# of Largest% of flaws % of flaws

2-parameter

2-parameter Technique

20POD of NDT Techniques

NDT Analysis – POD Results

• Basic POD analysis (2-parameter hit/miss= default)

0.00.20292.944.00.2790.203Turbo-MOI (Single)

0.00.20289.741.00.3400.246Turbo-MOI (Double)

0.0N/AN/A0.0N/AN/ATrecscan

0.0N/AN/A0.0N/AN/ASAM

0.3200.00.21442.912.00.2970.232RivetCheck (Single)

0.00.21446.212.8N/A0.457RivetCheck (Double)

0.2270.00.20473.718.00.2240.183Remote Field Eddy Current (Single)

0.00.20464.020.30.3680.254Remote Field Eddy Current (Double)

0.1400.00.15492.339.00.2180.162MWM (Single)

0.00.15494.735.00.2100.169MWM (Double)

0.00.23735.79.00.4440.295MOI (Single)

0.00.23735.98.00.3430.268MOI (Double)

flaw size(rate)Missed(>0.150)(All flaws)flaw sizeflaw size

detectedfalse callsFlawdetecteddetected90/950.9

AANC 4-parameter# of Largest% of flaws

% of flaws

2-parameter

2-parameter Technique

21

User must perform final analysis with Report Card as ‘guidance’

NDT Analysis – Report Card (Caveats)

• Data collected in 2003 – many changes likely in inspection sensor and platforms

– Snapshot still valid to identify potential ‘Emerging’ NDT

• POD can be calculated many ways– 2 parameter hit/miss = default method– Other POD methods examined: 4 parameter hit/miss, “a-hat” vs “a”, multi-flaw model,

variable angle model– Data will be made available in project database for further examination

• Screens/data available for “a-hat” versus “a” analysis (and S/N data)

• No cost data is presented; Must perform cost/benefit analysis (based on user)

• Capability versus reliability– Level IIIs versus inspectors

• External ultrasonic inspections hampered by inconsistent sealant

• Crack orientations affected some inspections– Very different from assumed orientations or previous studies

22Report card of NDT Techniques

NDT Analysis – Report Card

109915Turbo-Magneto Optical Imaging (T-MOI)

444-Trecscan

111-Structural Anomaly Mapping (SAM)

76714Rivet CheckTM

44318Remote Field Eddy Current

105820Medium Frequency Eddy Current

52619Meandering Winding Magnetometer (MWM)

5369MAUS Rotoscan

74612MAUS Rasterscan

109913Magneto Optical Imaging (MOI)

1010811LFEC (Sliding Probe)

877-High Frequency Linear Array UT

56717Giant Magnetoresistive Sensor (GMR)

567-Eddyscan

56616Digital Radiography

107810Detailed Visual (Internal)

2458C-scan Eddy Current

246-Conventional Film Radiography

5247Automated Couplant Ejection System (ACES)

41076Array Eddy Current

FieldabilitySpeedEase of UseSensitivity

23Categories provide user-preferences

NDT Analysis – Report Card (Summary)• MFEC results provided best sensitivity

– Internal versus external inspection

• External methods which provided best sensitivity: – Meandering Winding Magnetometer– Remote Field Eddy Current– Giant Magnetoresistive Sensor (GMR)– Digital Radiography– Rivet Check– High Frequency Linear Array Ultrasonic

• External methods which were the fastest: – LFEC Sliding Probe– Array Eddy Current– Magneto Optical Imaging (MOI)– Turbo-MOI

• External methods which are truly ‘Emerging’: – Meandering Winding Magnetometer– High Frequency Linear Array Ultrasonic– Remote Field Eddy Current– Giant Magnetoresistive Sensor (GMR)– Digital Radiography– Rivet Check

External methods which are ‘ready’: – LFEC Sliding Probe– MAUS– Magneto Optical Imaging (MOI)– Turbo-MOI

24

Summary• NDT was only one portion of project

– Field Inspections simulate actual in-service maintenance– Pre-teardown inspections conducted under climate-controlled environment

• No significant difference in Field/Pre-teardown results– Emerging NDT techniques analyzed

• Damage Characterization results impact NDT– Tunneling under clad– Sealant inconsistency, origin locations

• Comparisons of NDT Techniques performed– Protocol makes comparisons valid– 4 categories for rating: Sensitivity, Speed, Ease-of-use, “Fieldability”

• Probability of Detection analysis– Truth data from Damage Characterization matched to NDT results– 2-parameter, 4-parameter hit/miss, multi-flaw, variable angle methodologies– Signal analysis (“a-hat” versus “a”) also performed

• Project database, including all NDT scans, data– Available to the public soon!

replica-based crack inspection

Documents