new structural guidelines for dent allowables on … files/thurs/0130... · new structural...

New Structural Guidelines for Dent Allowables on Fuselages

Cornelis Guijt

USAF Academy Department of Engineering MechanicsCenter for Aircraft Structural Life Extension

Oklahoma City Air Logistics Center

CONTENTS

• Overview• Research objectives• Field survey• T.O. overview and dent definitions• Fatigue considerations• Static stability • Summary• Recommendations• Future work

OVERVIEW

• Dented fuselage structure• Heavy maintenance burden• Oklahoma City ALC has requested research effort to

reduce maintenance burden• Boeing Field survey• Tech Order (T.O.) guidelines, proprietary OEM information• Prevent unnecessary maintenance by developing better

guidelines based on research• Current allowables depend on zones on the aircraft:

– Some zones no dents allowed – > 3” away from stiffener– 0.03”- 0.25” depth limit

RESEARCH OBJECTIVES

• Are dents a fatigue concern?– Effect of re-forming?

• Do dents affect static stability? – Effect on compression loading– Effect on shear loading

• Provide experimental data for future analysis• Recommendations for ALC and T.O.

FIELD SURVEY

• Field survey of 27 aircraft– 1127 dents

• 664 unfilled• 298 filled• 156 repairs

– Maintenance– Stands– Hail/strikes

OVERVIEW of typical T.O. LIMITS

• Hand forming not allowed• Fuselage zones

– 2024-T3 skins• <3” stringer, <0.03”, rest <0.25”• Fill for aero

– 7075-T6 skins• Must be repaired

• Multiple dents in one bay must be repaired• Dent on stringer must be repaired

• Ring support for diameter control• Potential energy for depth

– Weight– Height

Dent forming

FATIGUE CONSIDERATIONS

• Unstiffened panels– 0.04” thick, 2024-T3 Clad– Variations:

• Bare• 0.06”

• Critical Fatigue location– Not in bottom of dent– Ring support at critical location– Dent behaves like an open hole– Secondary bending due to shift in

neutral axis

dt16

0

0.1

0.2

0.3

0.4

0.5

0.6

234000 236000 238000 240000 242000 244000 246000 248000

cycles

a (i

nch) skin

dentaverage

Load bypass around dent

Multiple Dents (analysis)

• Up to ½ dent diameter, negligible effect• At 1/8 dent diameter, Kt increases 1.6 times compared to

single dent• Aspect ratio of dent/gouges do affect Kt

• Limited dent depth effect• Reforming dents improves fatigue life• Life well above service life (est. <100,000 cycles)

Cycles to failure vs. dent depthFatigue loading 17 Ksi, R=0.05

0200,000400,000600,000800,000

1,000,0001,200,0001,400,0001,600,0001,800,000

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4Depth (in)

Cyc

les

to fa

ilure

Stif fener proximity allow able Single dent allow able

Standardized fatigue testing

Reformed dents

• Limited effect of dent depth • Use of ring support and sharp transition

yields conservative results • Reforming improves but does not

completely restore fatigue life – Reduce secondary bending– Cold working

• Worst dent still has sufficient fatigue life • Dents do affect fatigue, but do not

decrease the required fatigue life (~100,000 cycles)

Fatigue Conclusions

STATIC STABILITY; COMPRESSION

• Proper compression loading• Analysis of compression loaded panels• Stiffened panels

– 0.04” 2024-T3 skin– 8” stiffener spacing

• 1000 countersunk fasteners• 7075-T6 stiffeners

– 20” panel length/frame spacing• Single dent (2” and 5”), Multiple 2” dents

• Verify proper loading• Effect of panel width• Non-effective skin (avg. stress)

Mid-bay strains 3 (6)

-1000

-800

-600

-400

-200

0

200

400

600

800

0 2000 4000 6000 8000 10000 12000 14000

Load (lbs)

mic

ro s

train

gage 3-24gage 6-24gage 3-16gage 6-16gage 3-14gage 6-14avg 3&6

24”, 16”, 14”

8”

20”

Compression panel

• Relief from multiple dent limitation• Effect of dents close to stiffener• Close to effective width area

Effect of multiple dents

• Inter-rivet buckling• Local stiffener crippling• Skin wrinkling• Torsional instability

Typical compression behavior

FEM RESULTS: CompressionAnalysis by Jim Greer, CAStLE

• Failure loads of panels;

• No significant effect of dents• No effect of multiple dents• Dents seem to stiffen/strengthen the panel slightly

0.620.25-dents

0.318.6Large dent

0.318.3Pristine

KipKip

+/-avg

Max LoadPanel

Compression testing conclusions

STATIC STABILITY; SHEAR

• Bi-axial shear loading more complex• Shear:

– Tension in one direction, equal compression perpendicular to tension

– Tension: no stability concern– Compression: possibly same behavior as earlier in

the compression test • Single dent (2” and 5”), Multiple dents (2”)

Shear loading

• B-52 example

• Shear buckling apparent on static display/ground loading

• Shear buckles disappear in flight due to lift of fuselage

Large shear frame

20”

8”

20”

Shear conclusions

• Shear failure loads:

• No detrimental effect of dents– Slight increase of failure loads with dents

0.736.45-dents

0.434.6Dented

0.834.0Pristine

KipKip

+/-avg Max LoadPanel

Summary

• Dents do affect fatigue • Dents do not diminish the life under the required fatigue life

• Damage in the non-effective skin area will not affect panel strength in compression

• Most of skin is ineffective in shear/compression• Only damage in the effective width of the skin might possibly

affect panel strength• Very limited analysis needed; effective/non-effective skin

Recommendations

• Favorable T.O. recommendation: dents outside the effective width of a panel are not an issue for static stability, no maintenance actions required

• A typical transport fuselage structure was tested, these results should be applicable for similar types of shell structures/aircraft

FUTURE WORK

Dents in 7XXX aluminum – Especially regarding fatigue concerns

• Expand to configurations with more effective skin – Thicker skins

• Less likely to dent?• Larger effective areas

– Effect of dents inside effective skin• Analyze effective zones on aircraft• Evaluate alternative repair options for dents

– Reforming techniques– Traditional repairs/cut-outs

• Bonded repairs at critical positions around dent • Cross-cutting solution• Results already applied on 2 platforms• Return on investment 20 - 600

QUESTIONS?

The views expressed are those of the author and do not reflect the official policy or position of the US Air Force, Department

of Defense or the US Government

Aging Aircraft Repair Strategies Utilizing Cold Expansion Technologies Aging Aircraft Repair Strategies Utilizing Cold Expansion Technologies

by: Len ReidVice President,Research & Development

29 November – 1 December 2005

Memphis Tennessee

ObjectiveObjective

Present an overview of the various cold expansion processes to show how these may be utilized to effect economical repair strategies on aging military aircraft and reduce maintenance cost

Aging AircraftAging AircraftAverage USAF Fleet Age

8.9 9.2 9 9.911.1 12.2

13.1 13.914.7 15.6

16.7 17 17.3 17.618.9 20.6

2223.4

8.5

23.1 23.6 23.924.9

0

5

10

15

20

25

30

67 69 71 73 75 77 79 81 83 85 87 89 91 93 95 97 99 01 03 05 07 09 11

Projected

MXG Briefing; Gen (Select) Cameron

OO-ALC Nov 2005

OPR: Name & Organization

Maintenance Costs Growth (projected)Maintenance Costs Growth (projected)

2769.6

3582.6 3674.1 3600.4

4522.3 4453.54771.3

5105.8 5161.65425.9

5920.7

0

1000

2000

3000

4000

5000

6000

FY01

FY02

FY03

FY04

FY05

FY06

FY07

FY08

FY09

FY10

FY11$M FUNDING

MXG Briefing; Gen (Select) Cameron OO-ALC Nov 2005

What are the structural maintenance cost driversWhat are the structural maintenance cost drivers� Structural inspections� Fatigue crack abatement� Corrosion abatement� Lug durability/bushing migration� Fuel leaks� Component replacement� Major Structural Life Improvement

Programs (SLIP)

Maintenance cost elementsMaintenance cost elements

Today’s Maintenance Challenges Require New

Solutions

Today’s Maintenance Challenges Require New

Solutions

Structural Problems Solved By Residual Compressive Stress InitiativesStructural Problems Solved By Residual Compressive Stress Initiatives

Hole resizing, bushing retention, life and damage tolerance enhancement of lugs

Elimination of nut plate fatigue, fastener hole resizing

Blind access and improved sealing of blind threaded nuts and inserts

Structural fatigue life and damage tolerance enhancement

Improved structural modifications, design and repairs

Cold Expansion SolutionCold Expansion SolutionCold Expansion (cold working) induces a zone of residual compressive stress around and through a hole, typically extending radially at least one radius around hole.

Improves fatigue life and durability and damage tolerance of structure.

Hole is effectively “shielded,” reducing effective stress intensity factor and therefore the propagation fatigue cracks.

Hole is effectively “shielded,” reducing effective stress intensity factor and therefore the propagation fatigue cracks.

Radial expansion of bushings is also effective in inducing beneficial residual stresses in attachment lugs and fittings to enhance durability and fatigue life

cburns

click on above photo to view video clip

Extended Inspection Interval with Cold ExpansionExtended Inspection Interval with Cold Expansion

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

100 1000 10000 100000 1000000

Cycles (N)

Crac

k Le

ngth

(inc

h)

1 3A2A1A32

Critical Crack Length

Non Cold Worked

Cold Worked

Constant amplitude fatigue test

Material: 2024 T3

Hole Dia: 0.312 in

Stress: 25 ksi net, R-Ratio: 0.05

Environment: Ambient air

Cold expansion will often extend initial inspection threshold beyond the original economic life of the structure

Reduced Inspections Engine Pylon ExampleReduced Inspections Engine Pylon Example

Flight Cycles

ForceMate eliminates costly inspectionsForceMate eliminates costly inspections

� ForceMate bushings replaced shrink fit plus roller-swage bush� Rework time cut to hours rather than days� Most significant was extended inspection intervals

AH-1W COBRA Stub Wing Lug AttachmentAH-1W COBRA Stub Wing Lug Attachment� Production and overhaul retrofit to increase fatigue

life of wing from 1500 to 4000 hours� Use of ForceMate bushings realized

operating and support cost savings of $24.1 to $44.3 million

Fatigue Crack Abatement

Stress IntensityCausing CrackGrowth

Without Cold Expansion

With Cold Expansion

Crack TipPosition

Cold Expanded MaterialAround Hole

The large zone of residual compressive stress� Improves damage tolerance � Reduces stress intensity factor

range (∆K)� Retards or arrests crack growth

Reduction in Crack Tip Stress Intensity

Schematic RepresentationReport G. Clark, DSTO ARL, AUST.

1991 Fatigue of Engineering Materials

. 1 2

. 1 0

. 0 8

. 0 6

. 0 4

. 0 2

1 03 1 04 1 05 1 06. 0 0

Non-Cold Expanded

Cold Expanded

Cycles to Failure

Initi

al C

rack

Len

gth

(inch

) Initial Crack L

ength (mm

)

3 . 0 0

2 . 5 0

2 . 0 0

1 . 5 0

1 . 0 0

. 5 0

. 0 0

7075-T6 AlNet Stress = 30 ksiConstant AmplitudeR = 0.051/4" dia. holesPetrak & StewartEFM 1973

Durability and Damage Tolerance (DADTA) BenefitDurability and Damage Tolerance (DADTA) Benefit

� USAF DADTA* Approach

�Assume 0.050-inch initial flaw

When hole cold expansion used�Allow smaller initial flaw

�Typically 0.005 inch

Very conservative analysis method

*MIL-A-83444, MIL-STD-1530

Conservatism of USAF DADTA ApproachConservatism of USAF DADTA Approach

Constant Amplitude Fatigue TestMaterial: 2024-T3Hole Diameter: 0.312 inWidth: 2.35 inThickness: 0.199 inStress: 25 ksi net, R-Ratio: 0.05Freq.: 10 HzEnvironment: Ambient Air

Constant Amplitude Fatigue TestMaterial: 2024-T3Hole Diameter: 0.312 inWidth: 2.35 inThickness: 0.199 inStress: 25 ksi net, R-Ratio: 0.05Freq.: 10 HzEnvironment: Ambient Air

Cold Expanded actual life0.005 initial flaw670,000 cycles

Cold Expanded actual life0.005 initial flaw670,000 cycles

Non Cold Expanded actual life0.005 initial flaw40,000 cycles


InitialCrackLength(inch)

MIL-A-83444 requirements (Non Cold Expansion)

Min allowable crack size fordesign/analysis incorporating Cold Expansion

10000 100000 1000000

0.000

0.025

0.050

0.075

0.100

0.125

Cycles to Failure

Non ColdExpanded

Cold Expanded



Damage Tolerance Benefits; Fighter Aircraft StudyDamage Tolerance Benefits; Fighter Aircraft Study

Ref: Rich, D. L. and Impellizzeri, L. F., “Fatigue Analysis of Cold-Worked and Interference Fit Fastener Holes,”ASTM STP 637 1977

No Cold Expansion Split Sleeve Cold Expansion

Repair of F-16 Fuselage Access Panel Riveted Nut PlatesRepair of F-16 Fuselage Access Panel Riveted Nut Plates

Upper Fuselage Skin Access Panels

Fighter Doors Program

Fatigue Test Results

0

5000

10000

15000

20000

25000

30000

35000

40000

Baseline RivetedNutplate

ForceTec Repair ForceTec in new aluminum skin

Equi

vale

nt F

light

Hou

rs (E

FH)

Sealed ForceTec Rivetless Nutplate Installation

Home Plate Specimen

cburns

click on photo to left to view video clip

Corrosion AbatementCorrosion Abatement

� Cold expansion does not stop corrosion however, it can mitigate the damage in holes if not too severe by allowing removal of damage and installation of oversize fasteners

� ForceMate bushings can repair/resize the hole to nominal size

cburns

click on photo to view video clip

ForceMate SolutionWing Weapon PylonForceMate SolutionWing Weapon Pylon

� FTI ForceMate solves bushing corrosion problems� Large 5 ½� bushing� �BlueCoated� for

anti-fretting

F-18 Wing Fold Transmission ForceMate Repair SolutionF-18 Wing Fold Transmission ForceMate Repair Solution

� Wing fold transmission repair using FTI ForceMate bushings

� ForceMate allows repair of corrosion problems without removal of �slices�

ForceMate Bushing to Repair Rainbow Fitting CorrosionForceMate Bushing to Repair Rainbow Fitting Corrosion

Most common cause of fitting replacement is corrosion. USAF ForceMate allows fittings to be repaired on aircraft versus being scrapped

Cost of tooling recovered on first aircraft completed

Graph represents cost per aircraft difference

Fittingreplacement

ForceMate

C-130 Cargo Ramp CorrosionC-130 Cargo Ramp Corrosion

� Problem: Corrosion in attachment of sloping longeron typically requires longeron replacement

� Solution: Corrosion in holes removed and holes restored using expanded bushings

� Savings exceed $75,000 per fitting restored

Corroded fitting

Repaired fitting

Lug Durability/Bushing migrationLug Durability/Bushing migration

� ForceMate® Induces beneficial residual stresses in parent material� Increased fatigue life

& damage tolerance� Consistent high

interference fit � Provides greater

resistance to rotation, fretting and push-out

Typical lug failure from fretting induced fatigue

100,000cycles @ 17 KSI

Same bushing installed using ForceMate

1,000,000 cycles @ 21 KSI

Typical Fatigue Life ComparisonShrink Fit and ForceMate Bushings

137.9

103.4MaximumNet SectionStress Level

(MPa) 68.9

34.5

2 2 24 4 46 6 68 8 8105 106 107

Shrink Fit

X

XX

X

XXXXX

XXXX

XXX

X

X X XX

XX

XXX

Test Specimen7075-T651

Load Conditions:Constant amplitude10 HzR = .05Beryllium Copper Bushings

ForceMate FailureShrink Fit FailureNo Failure

Champoux & LandyASTM STP 9271987

cburns

click on above photo to view video clip

ForceMate Bushing Damage ToleranceForceMate Bushing Damage Tolerance

ForceMate crack growth life improvement - 20:1

1 10 100 1000 10000 100000 1000000 0.00

2.00

4.00

6.00

8.00

10.00

0.000

0.079

0.157

0.236

0.315

0.394

Crack Length (inch)

Crack Length (mm)

Cycles

100% Load Transfer Axial Fatigue Technology Inc. Material: Titanium Maximum Load: 90 kN (20 Kip), R-Ratio: +0.10 Frequency: 20 Hz, Environment: Ambient Lab Air Note: All specimens were corner notched and precracked prior to testing.

Shrink Fit

ForceMate

�17-4 stainless steel ForceMate bushings

�28.5 mm (1-1/8 inch) pin diameter

�Titanium lug specimens�Pre-cracked with corner flaw in hole�Two lugs with outer edge cracks

�17-4 stainless steel ForceMate bushings

�28.5 mm (1-1/8 inch) pin diameter

�Titanium lug specimens�Pre-cracked with corner flaw in hole�Two lugs with outer edge cracks

� Migrating bushings� Sealant bond breaks� Corrosion results� Old process was recurring

inspections, and fitting replacement when beyond limits

� ForceMate repair� Migration is eliminated� Sealant bond remains� Corrosion in lug bore is

eliminated

Horizontal Tail Attachment Lugs Bushing migration and subsequent corrosionHorizontal Tail Attachment Lugs Bushing migration and subsequent corrosion

Torque VibrationPushout Fatigue

Comparative Performance to Shrink-Fit BushingsComparative Performance to Shrink-Fit Bushings

3933

22107

0

500010000150002000025000

Shrink-Fit ForceMate

107

1090

0

500

1000

1500

Shrink-Fit ForceMate

15150

0

50

100

150

Shrink-Fit(Max.)

ForceMate(Min.)

10 : 1 Improvement10.2 : 1 Improvement 5.6 : 1 Improvement

> 10:1 at 25% increase in stress level

Engine Strut ForceMate ModificationEngine Strut ForceMate Modification

� B-52 fleet has been modified� KC-135 modification in-work

�Modification eliminates fatigue inspection of strut attach locations

C-130 Truss Mount Cone Bushing Hole and Sway Brace AttachmentC-130 Truss Mount Cone Bushing Hole and Sway Brace Attachment

ForceMate Bushings saved 150 maintenance man-hours per truss

Component ReplacementUSAF SavingsComponent ReplacementUSAF Savings

� C-130 Rainbow Fitting� F-16

� Upper Fuselage skins� Pylon Rib� Falcon Up � bulkhead replacement

� KC-135 Wing skin replacement� Vertical tail LE spar

F-16 Falcon UPF-16 Falcon UP

� F-16 Wing carry through bulkheads

� Fatigue cracks found in the durability test, and in the field

� Next Best Alternative �replace bulkheads

� Falcon UP = Split Sleeve Cold Expansion to size, less than 10% of the cost

Projected cost

ReplaceBulkheadsFalcon Up

Cold expansion of non-circular holes

KC-135 – ECP-484KC-135 – ECP-484

� Lower wing skin cracking in the outboard wings

� Option: Skin or wing replacement

� Solution � Split Sleeve Cold Expansion to size �Less than 1/4 of the cost to implement

Projected cost

ReplaceskinsECP-484

ForceTec Riveted Nut Plate ReplacementForceTec Riveted Nut Plate Replacement

Leading Edge attach holes damaged and ovalized

� Need to install new LE� Current method would

require over sizing holes to match damaged holes in LE spar

� ForceTec repairs and resizes holes in one operation

� Solves Edge Margin concerns in repair

� In 2003 78 USAF aircraft in depot for SLIP mods

� 862 extra days unplanned downtime due to fuel leaks from wing fasteners � (ten days per aircraft)

� In addition, operators report fuel leaking as top driver for maintenance effort and reduced mission capability rates.

F-16 Wing Fuel Leak Problems

The Solution - TukLocThe Solution - TukLoc� Advanced blind fastening

system with the advantages of a high interference fit from FTI�s cold-expansion technology.

� Direct replacement for NAS1734 nuts

�Sample 8 jets 16 months prior to TukLoc

�66 leak occurrences�~3 days downtime/jet

�Includes sealant cure time

�238.4 maintenance man-hours

�Since TukLoc – zero leaks reported

Joe Smith, Analyst

OO-ALC 508th Fighter Support Wing

Results of Rework

cburns

click on photo to view video clip

Major Structural Life improvement programsMajor Structural Life improvement programs� Wing structural life extensions

� KC-135 Wing (ECP-484)� T-38 Wing enhancement� A-10 Hog-up� USN P-3 Orion� F-4

T-38 Lower wing skin Countersink Cold Expansion of 200 fatigue critical holesForceMate repair of D-panel existing riveted nut plate holes

Life extension mod to P-3 FleetLife extension mod to P-3 Fleet

Fatigue cracks initiated from the fastener and nutplate satellite rivet holes in the under wing fillet fairing.

Wing life extended by cold working satellite rivet holes and installing ForceTec Rivetless Nutplates.

Alternative is inner wing box replacement

Wing FSFT experienced cracking at 16,785 test hours

Repair Test Demonstrated Life84,000 Hours

Spiraling Maintenance Costs can be constrained or reducedSpiraling Maintenance Costs can be constrained or reduced

New Challenges require New SolutionsNew Challenges require New Solutions

Exploit the advantages of innovative technologiesExploit the advantages of innovative technologies

SummarySummary� Innovative thinking needed to reduce the total

operating cost of our aging fleet� Don�t accept �that�s the way things are.�� New technology is available to reduce costs

� In all cases, the applications addressed in this briefing were paid for in reduced cost of the maintenance activity.

“Cold expansion technologies can be directly credited with keeping the fleet of military aircraft flying and saving millions of dollars in the process”

Dr. Jack Lincoln

Questions?Questions?

Thank You!Thank You!

Ogden Air Logistics Center

BE AMERICA’S BEST

Development of an Inspectable Repair for A-10 Fuel Vent Hole

Cracking in Integrally Stiffened Wing Skins and Extruded Spar

Caps

1Lt Larry WareA-10 System Program Office

[email protected]

Aircraft Structural Integrity Conference

01 Dec 2005

BE AMERICA’S BEST

O G D E N A I R L O G I S T I C S C E N T E R

Co-Authors!Ms. Kristi Xidas- A-10 Test Engineer, NGC!Dr. Paul Clark- A-10 Analysis Lead Engineer, USAF!Dr. Mark Thomsen- A-10 ASIP Manager, USAF

!Special Acknowledgements! Ken Grube- A-10 Program Manager, NGC! Ken Eaton- Regional Manager, FTI

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Overview!A-10 Wing Construction!Repair Development!Validation Process!Repair Process!Cost Benefit!Other Applications

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A-10 Wing!The A-10 has met its original service requirement

and must now serve at least a second!A-10 wing consists of 3 sections

! Wing Center Panel (WCP)! 2 Wing Outer Panels (WOP)

!Wing Center Panel! Approximately 18 ft long! Integral fuel cell! 2 configurations- thin lower skin and thick lower skin! Service Life Extension Program (SLEP 1)

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Wing Construction!WCP Upper Skin

! 2 integrally stiffened machined 7075-T76511 extruded planks

! 80 vent holes per wing !Mid-Spar Upper Cap

! T-section 7075-T76511 extrusion! 8 vent holes per wing

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Upper Cover Cracking!Cracks form on upper and lower sides of holes!Average of 16 holes per aircraft are found cracked!Maximum of 64 holes found cracked on one aircraft

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Original Repair! 2 nested back to back L-angles! 18 new fastener holes! 9.5 hours labor and fabrication

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Damage Evolution!Residual Tensile Stresses around the open hole

were the root cause of the cracking! Initial compressive loading produces localized

yielding around vent holes! Compression cycles combined with residual stresses

on top and bottom of holes develop cracks

Local Compressive

Yielding

Residual Tensile Stress

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Damage Evolution

Figure 6-2 Northrop Grumman Report SA220R0438

!Compression dominated spectrum which also includes some tension loads due to negative bending increases stresses amplitude

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Improved Repair!New Repair Concept

! Install steel bushing or plug to prop hole open and still transfer compressive loads

•Bushing Migration?????•Hole Fit????

! Fatigue Technology Inc. (FTI) ForceMate Bushing Installation

•Mandrel to expands bushing in the hole•Similar to BushLoc•No split sleeve required•Restrains hole from elongating•Better hole fill improves durability

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Validation Testing! Installation test

!A-10 spectrum fatigue test! Flat plate coupon test! Semi Component test

!Static Testing! Semi Component test

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Flat Plate Coupon Testing!Flat plate coupon testing

! 3- original .375” baseline holes! 3- 3/8” cold worked holes! 3- .500” ForceMate bushed holes

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Flat Plate Coupon TestingFlat Plate Coupon- 3/8" Cold Worked Hole

0.000

0.020

0.040

0.060

0.080

0.100

0.120

0 5000 10000 15000 20000

Effective Flight Hours

Cra

ck L

engt

h, in

ches

Front LeftFront RightBack LeftBack Right

Flat Plate Coupon 3/8" Hole

0.000

0.020

0.040

0.060

0.080

0.100

0.120

0 5000 10000 15000 20000Effective Flight Hours

Cra

ck L

engt

h, in

ches

Front LeftFront RightBack LeftBack Right

•Cold Worked 3/8” hole showed similar crack growth to Original 3/8” hole configuration. No benefit in compression dominated spectrum

•.500” ForceMate coupon showed no cracks

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Semi Component Testing!A-10 loads spectrum testing

! Semi-component coupons•Components came from actual aircraft wing with 5832 flt hrs•3- original .375” baseline holes•3- .500” ForceMate bushed holes•3- 1.000” ForceMate bushed holes

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Semi Component TestingSemi Component Coupon- 3/8" Hole

0

0.02

0.04

0.06

0.08

0.1

0.12

0 5000 10000 15000 20000

Effective Flight Hours

Cra

ck L

engt

h, in

ches

Left Side 1Left Side 2Right Side 1Right Side 2

•Solid lines indicate the cracks that are on the side of the hole closest to the skin

•Dashed lines indicate the cracks that are on the side of the hole furthest from the skin

•No cracks were found in either the .500” or the 1.00” ForceMatebushed holes

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Semi Component Static Testing Semi-Component Static Compression Test

0100002000030000400005000060000

700008000090000

100000110000120000

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14Deflection - compression - inch

Load

- co

mpr

essi

on -

lbf

3/8" Baseline Hole1/2" ForceMate Bushing1" ForceMate Bushing

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Testing Summary!Flat Plate Coupons

! Original 3/8” hole- 25,916 EFH=0.122” crack! Cold Worked 3/8” hole- 18,480 EFH=0.112” crack! ½” ForceMate Bushing- 39,990 EFH= no crack

!Semi-Component Coupons! Original 3/8” hole- 24,480 EFH=0.115” crack! ½” ForceMate Bushing- 16,800 EFH= no crack! 1” ForceMate Bushing- 22,254 EFH= no crack! All cracks verified at 35x magnification

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Testing Summary con’t!Wing Fatigue Test showed similar cracking to semi

component tests! Fleet cracking has shown larger crack sizes

!Semi Component Static Compression Test! Original 3/8” hole configuration

•Baseline data! ½” ForceMate Bushing Installation

•No loss in crippling strength! 1” ForceMate Bushing Installation

•No loss in crippling strength

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Repair Process

! Tooling Requirements•Pre-Reamer•Starting Reamer•Mandrel•Offset Puller•Hand Pump•Bushing

! Repair Steps1.Ream hole with pre-reamer2.Ream hole with starting reamer3.NDI to ensure no cracks4.Install bushing using offset puller and hand pump

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Installation

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Inspection Capability!Repair can still be inspected by the current

inspection method with the bushing installed!Due to the success of the tests, inspection interval

is now 8000 hrs for the A-10

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$ Cost Benefit $!Original Repair

! 9.5 hours labor = $937.00! Repair components = $48.00! Total cost = $985.00

!ForceMate Bushing Repair! 1 hour labor = $98.00! Bushing cost = $38.00! Total cost = $136.00

!Total Savings! Per repair- $849.00! Per Aircraft (avg. 16 repairs)- $13584.00! Remaining SLEP1 aircraft- $3 MIL

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Other Applications!Lower Skin Integral Stringer Fuel Transfer Holes!Pylon Stud Holes!Wing Attach Fitting Holes!Upper Spar Cap Fastener Holes

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Summary!A-10 Wing Construction!Repair Development!Validation Process!Repair Process!Cost Benefit!Other Applications

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Questions ?

[email protected]

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