recent developments for improve and assess repair ... · pdf filevienna, 19-21 june 2006...

Vienna, 19-21 June 2006 Presented by: George Lampeas, Univ. Patras, Greece

Recent developments to Improve and Assess Repair Recent developments to Improve and Assess Repair Capability of Aircraft StructuresCapability of Aircraft Structures

G. Lampeas, Univ. Patras - Greece

C. Meyer, Airbus - France

Recent Developments for Improve and Assess Repair Capability of Aircraft Structures



CC

ContentsQ 0. Main objectives of the European project IARCASQ 1. Fuselage Riveted Skin and Joint RepairsQ 2. Fuselage Bonded Skin RepairQ 3. Fuselage Welded Stringer RepairQ 4. Fuselage dentsQ 5. Wing box repairs and allowable DamageQ 6. Friction Stir Welding Repairs and Portable FSW Crack repairsQ 7. LE Bird Impact RepairsQ 8. Validation Expérimental TestsQ 9. Conclusions



0. IARCAS Objectives 0. IARCAS Objectives and brief overviewand brief overview

(Contract G4RD(Contract G4RD--CT2000CT2000--00401)00401)Overall Objective

To improve repair efficiency of aircraft metallic structures in order to realise a significant reduction of operational and maintenance costs

Quantitative Objectives1. To reduce the downtime of the aircraft by 25%

2. To increase inspection intervals by more than 20% for specific repairs

Technical Targets1. To develop a better approach for the sizing criteria governing the structural strength

of repairs and allowable damages

2. To develop and implement repair procedures for new technologies and materials

3. To increase the fatigue life using new repair techniques, materials, processes



0. IARCAS Overview Presentation 0. IARCAS Overview Presentation Work Work BreakdowmBreakdowm StructureStructure

WP3Improvement / Development of design principles for

repairs and allowable damages

WP1Review and classification of current repair techniques and data

WP2Methods development and

improvement

WP4Testing and validation of calculation methods

and improved repair principles

WP5Costs and benefits of improved repair principles

Divisione AeronauticaDivisione AeronauticaDivisione Aeronautica



1. Fuselage Riveted Skin and Joint Repairs1. Fuselage Riveted Skin and Joint Repairs

Main results : a) Improvement of calculation

methodologies by developing detailed non-linear fully parametric FEM

b) Accurate determination of inspection threshold and inspection intervals

c) In certain cases possibilities to increase them as compared to standard methods were identified



1. Riveted doubler repairs:1. Riveted doubler repairs:The development of a The development of a detailed nondetailed non--linear fully parametric FEM allows to:linear fully parametric FEM allows to:

•• Perform Fatigue life assessment of Perform Fatigue life assessment of external and internal skin repairsexternal and internal skin repairs

•• Investigate the effect of different Investigate the effect of different repair parameters on the IQF (repair parameters on the IQF (Effect of Effect of doubler thickness and material, rivet doubler thickness and material, rivet diameter, longitudinal and transverse diameter, longitudinal and transverse rivet pitch, corner fastener location and rivet pitch, corner fastener location and biaxial loading)biaxial loading)

•• Calculate the stress field, Calculate the stress field, SIFsSIFs, crack , crack growth and residual strengthgrowth and residual strength

•• Compare Results with SRMCompare Results with SRM

Example: Example: Section Section 11/12 of 11/12 of A330A330



1. Riveted doubler repairs 1. Riveted doubler repairs -- Effect of rivet diameterEffect of rivet diameterSkin thickness 1.6 mm ; doubler thickness 1.8 mm ;

wl=wt=20 mm

100

105

110

115

120

125

SRM : 4&4,8 mm only 4 mm only 4,8 mm

IQF

• 4 & 4.8 repair performance is similar than a repair with only 4 mm rivets since the first row is critical• The load transferred in the critical fastener is lower with 4 mm rivet than for 4.8 mm rivet. But, the Kt due to the load transferred is higher (smallest hole)• With 4.8 mm rivets only, fatigue performance is increased by 12% compared to the baseline (factor 1.7 on fatigue life)



IQF=125 IQF=131 IQF=131Increase by 4.5% on IQF, i.e. 20% on fatigue life in the present case

• Life improvement factor by 1.1 to 1.2 may be achieved without any other modification (no additional fastener)

1. Riveted doubler repairs 1. Riveted doubler repairs --Effect of corner fastener location under Effect of corner fastener location under uniaxialuniaxial loadingloading



1. Riveted doubler repairs: Effect of cut1. Riveted doubler repairs: Effect of cut--out and doubler out and doubler shapes on fatigue for small skin repairsshapes on fatigue for small skin repairs

Rectangular cut-out with rectangular doubler (SRM) Circular cut-out with rectangular doublerImproved corner fastener location (R,θ) = (20;45) Improved corner fastener location (R,θ) = (20;45)

IQF 137 (6% improvement) 139

• No impact since the critical site is the corner fastener and not the cut-out radius



1. Riveted doubler repairs:1. Riveted doubler repairs: Effect of doubler thicknessEffect of doubler thickness3 Fastener rows

100

110

120

130

140

1,00 1,10 1,20 1,30 1,40 1,50 1,60

Doubler thickness / skin thickness

IQF

Epaisseur peau 1,2Epaisseur peau 1,4Epaisseur peau 1,6Epaisseur peau 1,8

• The higher the td/ts ratio, the lower the fatigue performance due to eccentricity• No improvement is possible by increasing doubler thickness for 2 and 3 fastener rows repairs

SRM values



Material 2024-T3–thickness 1,6 mm, containing a circular central cut-out (diameter 40 mm) is reinforced on external side with rivetedaluminium doubler (material 2024 T3 CLAD – diameter 114 mm -thickness 1,6 mm - Countersunk rivets NAS1097 DD, 4mm ).

1. Riveted doubler repairs : Circular riveted repairs1. Riveted doubler repairs : Circular riveted repairs

Calculated fatigue life : 250 000Test result : 340 909 FC

- Conservative FEM- Low Margin : 7% on allow. stress



1. Riveted doubler repairs: Crack Propagation in Skin repairs1. Riveted doubler repairs: Crack Propagation in Skin repairs

•• The nonThe non--linear FEM model may be used for Inspection Interval linear FEM model may be used for Inspection Interval calculationcalculation

Cracking scenario considered

Coupon 1 : 3 rivet rows repairs under 60 MPa

0

10

20

30

40

50

60

70

80

0 20000 40000 60000 80000 100000 120000 140000

Number of cycles

Hal

f cra

ck le

ngth

Analytical

Numerical - upper crack

Numerical - Lower crack

Test - Upper crack

Test lower crack



ÁÁ Fatigue life and residual Fatigue life and residual strength strength

1. Riveted doubler repairs:1. Riveted doubler repairs: Coupons testsCoupons tests

Al = Aluminium

Gl = Glare

CW = Cold Worked

FP = Flap Peened

IF = Interference Fit



2. Fuselage bonded skin repair2. Fuselage bonded skin repair

Main results : a) The hot bonded repair solutions demonstrated to be more

efficient in terms of longer fatigue lives than the standard riveted repair solution (SRM repair). Easy to perform surface pre-treatment methods have been introduced and shown fatigue performance similar to production type pre-treatment methods.

b) Development of calculation methodologies that were not available within commercial aviation up till now: Progressive Damage Models based on detailed fully parametric FE models and the StressCheck code

c) Improved bonded skin repairs by optimization of material and geometrical repair parameters



0

1

2

3

4

5

0 20 40 60 80Half patch width (mm)

SIF

(MP

a m

^1/2

)

[10]maximumminimum

0

1

2

3

4

5

6

0 20 40 60 80Half patch height (mm)

SIF

(MP

a m

^1/2

)

Effect of patch widthEffect of patch width

Effect of patch heightEffect of patch height

Repair by bonded double-sided composite patch (CFRP or Boron-epoxy)

2. Fuselage bonded skin repair 2. Fuselage bonded skin repair –– SIF calculation of CenterSIF calculation of Center--cracked plate cracked plate



Cracked hole (D=5mm, a=2.5mm, x=40mm)

0

5

10

15

20

25

30

35

0 50 100 150 200 250 300

Applied stress (MPa)SI

F (M

Pa m

^0.5

)

Patch (degradation)PatchNo patch

Effect of patching on SIF

2. Fuselage bonded skin repair 2. Fuselage bonded skin repair -- SIF evaluation of an SIF evaluation of an Open hole cracked plate repaired by a singleOpen hole cracked plate repaired by a single--sided sided bonded compositebonded composite patchpatch



To avoid patch debonding a tapered edge is used.

150 MPatapered non-tapered

250 MPatapered non-tapered

Patch debonding for 15 mm crack length

2. Fuselage bonded skin 2. Fuselage bonded skin repair repair –– Tapered edge Tapered edge patchpatch



0

1

2

3

4

0 100 200 300 400 500

Applied stress (MPa)

Dis

plac

emen

t (m

m)

one sideboth sidesboth sides + pinboth sides el-pl

2. Fuselage bonded skin repair 2. Fuselage bonded skin repair –– Repaired lapRepaired lap--jointjoint



crack x1

x2

CFRP patch

2024-T3 plate

x3x3

LOADING GRIPS

LOADING GRIPS

single-sidedrepair

back-to-backrepair

2. Fuselage bonded skin repair 2. Fuselage bonded skin repair –– Comparison of Comparison of single sided and backsingle sided and back--toto--back repairback repair

0

5

10

15

20

25

30

35

40

45

10 20 30 60

a [mm]

r.m.s

. SIF

[MPa

.m1/

2 ]

single-sided: S = 10MPa

single-sided: S = 100MPa

back-to-back: S = 10MPa

back-to-back: S = 100MPa



ÁÁ Fatigue life and residual strength Fatigue life and residual strength 2. Fuselage bonded skin repairs 2. Fuselage bonded skin repairs -- Coupons testsCoupons tests

HB = Hot bonded

(FM-73M)

CB = Cold bonded

R = “In-Field”Repair

P = Production



3. Fuselage 3. Fuselage welded stringer repairs welded stringer repairs Main results:

Q Specific repair design for welded panels are not required. Same coupling, fasteners, sealant used for riveted repairs are enough, at least for single aisle aircraft family. Q This has been demonstrated both for fatigue and static loading under compression on representative full-scale tests. In particular, repair should not be a No-Go item for potential laser beam welding applications in some upper fuselage shells.



Detailed Finite Element Model Evaluation of the effect of the stringer stringer thickness / skin thickness / cutting angle

6874 72

6366

60

78

70

55

70

0

10

20

30

40

50

60

70

80

90

2,5 / 2,2 / 0° 2,5 / 2,2 / 45° 2 / 2,2 / 0° 2,5 / 1,2 / 0° 2,5 / 1,2 / 45°

Type of stringer run-out (stringer thickness / skin thickness / cutting angle)IQ

F

IQF exp

IQF cal

3. Fuselage 3. Fuselage welded stringer repairs welded stringer repairs -- Prediction of Prediction of crack growth behaviour along the welding linecrack growth behaviour along the welding line



50

55

60

65

70

75

80

85

90

95

100

10000 100000 1000000Nb of cycles up to failure

Max

imum

gro

ss s

tress

2,5/2,2/0 2,5/2,2/45 2/2,2/0 2,5/1,2/45 2,5/1,2/0

3. Fuselage 3. Fuselage welded stringer repairs welded stringer repairs –– Coupons testsCoupons testsID 2



QQ A non-linear elastoplastic fully parametric tool for the simulation of the denting process

Q Allowable dent limits were extended, using FE fatigue analysis and tests from coupons to full-scale tests.

Q The influence of the applied dent depth on flat and curved panel static failure load was identified.

4. Fuselage dents 4. Fuselage dents -- Main results:

Dents Classification

Example: AExample: A--320 rear door area320 rear door area



4. Fuselage dents 4. Fuselage dents -- Simulation of denting process of Simulation of denting process of Riveted, Bonded or Welded stiffened fuselage panels for Riveted, Bonded or Welded stiffened fuselage panels for RS calculationRS calculation

A380 Welded Panel FE Denting Predictions

0

5

10

15

20

25

Time (s)D

ispl

acem

ent (

mm

)

Max Panel DispMeasured Panel Disp



Step 1 : dent modelling

tool indenting

Springback effect→ residual depth

Step 2 : application of the fatigue load

Pmax Pmin

Step 3 : fatigue criterion computation

grey : fatigue life>106

fatigue life ≈ 300 000

4. Fuselage dents 4. Fuselage dents -- FE fatigue analysis FE fatigue analysis



0,E+00

1,E+05

2,E+05

3,E+05

4,E+05

cycl

es n

umbe

rFE calculations

tests

6056T78th=1,2

Pmax=140MPares.dent=6mm

6056T78th=1,8

Pmax=140MPares.dent=7,9mm

6056T78th=2,2


2524th=1,8


2524th=3,5


Evaluation of Evaluation of thethe allowableallowable dents dents predictionprediction approachapproach(FE (FE calculationscalculations andand fatigue fatigue criterioncriterion ofof 300000fc 300000fc ofof maintenance maintenance freefree lifelife))




Č The single aisle SRM dent diagram was extended compared to today’s maintenance free allowable dent limits

Č Extension to other applications (higher/lower skin thickness, lower indenter diameter)

1

2

3

4

5

6

1




5. Wing box repairs and allowable damage5. Wing box repairs and allowable damagei) Enhanced stopi) Enhanced stop--drill techniques and cold expansion/interference fit drill techniques and cold expansion/interference fit

repair of cracks at fastener holesrepair of cracks at fastener holesiIiI) Surface blends) Surface blends

Q Identification of the main parameters affecting the refurbishment of cold worked holes, including the effects of re-cold expansion and the enhanced stop-drill repairs Q An experimental and modelling activity has demonstrated the capacity to permit further blending of minor wing skin damage than is currently allowed by the Structural Repair Manual (SRM).

Central hole

M12 bolts for end fittings

Stop drillrepair

200mm

400m

m

2a

d D L

CENTRALHOLE

STOP-DRILLHOLE

STOP-DRILLHOLE

ORIGINALCRACK

ORIGINALCRACK

SECONDARYCRACK

SECONDARYCRACK

2b

Main results:



5. Existing wing box repairs and allowable damage5. Existing wing box repairs and allowable damagei) Enhanced stopi) Enhanced stop--drill techniquesdrill techniques

0

10

20

30

40

50

60

70

0 50000 100000 150000 200000 250000

Cycles

a(m

m)

XA19709 (No stop-drill)XA19710 (open hole, no CX)XA19710 stop-drill holeXA19711 (open hole, CX)XA19711 stop-drill hole

Q Max net σ = 100MPa, R = 0.1, constant amplitude loading

Q SD holes were drilled at both crack tips when a = 10mm.

Q SD holes were centred at t ahead of crack tip (t= test piece thickness = 6.7mm).

Q CX measured as 3.4%.

Q SD increased Nf¸ 26% with no

CX¸ 300% with CX



5. Wing box repairs and allowable damage 5. Wing box repairs and allowable damage ii) Surface blendsii) Surface blends

W=200 mm

2.322.32 2.29

2.872.57

0

0.5

1

1.5

2

2.5

3

3.5

0 1 2 3 4 5 6 7 8 9 10Distance from front face of the plate (mm)

Stre

ss F

act

Plain side Indent side

• Quantification of stress field change within the blend areaQuantification of stress field change within the blend area•• Determination of allowable blend limitsDetermination of allowable blend limits.



• Blend effects on

crack propagation

“Type A”200 mm, Centre Hole

No Blend“Type B”

200 mm, Offset HoleOffset Blend

“Type F”400 mm, Offset HoleOverlapping Blend

“Type E”400 mm, Offset Hole

No Blend

“Type D”200 mm, Offset HoleOverlapping Blend

“Type C”200 mm, Centre Hole

Centre Blend

Blend / hole geometry

Crack length, 200 mm-wide specimens

0

5

10

15

20

25

30

35

40

45

0 2000 4000 6000 8000 10000 12000Cycles

mm

2144Type A

2152Type C

A FRONT A REAR

B FRONT B REAR2147

Type B2156Type D

A 30 mm diameter blend 19.05 mm away from a 6.35 mm diameter hole

0.5

0.6

0.7

0.8

0.9

1

1.1

0 5 10 15 20 25

Distance fron hole centre (mm)

CPL

_Ble

nded

/CPL

_Bas

elin

e

Test Analysis

A 30 mm diameter blend partly overlap with a 6.35 mm diameter hole

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0 2 4 6 8 10 12

Dis tance from hole centre (m m )

CPL

_Ble

nded

/CPL

_Bas

elin

e

Analysis

Test

5. Wing box repairs and allowable damage 5. Wing box repairs and allowable damage ii) Surface blendsii) Surface blends



Portable FSW Machine

6. Friction Stir Welding Wing Box Repairs and 6. Friction Stir Welding Wing Box Repairs and Portable FSW Crack repairsPortable FSW Crack repairs

Main resultsQ Development of reliable repair schemes for welded wing structures and tools to predict the fatigue performance Q Coupon test results for bolted joints with fastener holes through weld material have given better fatigue lives than predicted Q Bonded composite patch repair has been shown to provide a promising repair technique for fatigue cracks in FSW structure Q Development of Portable Friction Stir Welding Machine applicable to leading edge skin crack repair



7. LE Repairs against bird impact 7. LE Repairs against bird impact -- Main results QQ Good predictions of bird impact repaired structures response. Q Improvements of the repair design for structures without ribs in the nose area have been performed.



• The numerical simulation methodology requires detailed modelling of the repair surface including:the repair patchthe backing platesthe rivet rows with suitable rivet failure modelscontact surfaces

LE geometryRepair mesh and Al repair patch (riveted)

Al backing plates and repair patch(riveted)

Rivet rows (non-linear spring elements)

Detailed FE model of a LE with repair and bird model

7. LE Repairs against bird impact 7. LE Repairs against bird impact –– FE model developmentFE model development



Test specification

Measured value

Angle

Speed

Bird weight

37°

180 m/s ± 7

1.814 kg

37°

173.8 m/s

1.814 kg

LE Structure

Nose diameter 393 mm curvature Al 2024 T42 2 mm thick Material

NAS1921C05 Fasteners 2 x 3 fastener rows rows

Bird impact with respect to the flight direction

LE and test rig for the bird impact test

7. LE Repairs against bird impact 7. LE Repairs against bird impact –– Experimental testsExperimental tests



Transversal patch breaking all along the internal 3rd rivet row.

The breaking extend over the repair limits

Transversal top skin breaking along track rib flange. Parallel to the first one. Track rib flange not damaged

Track rib web

broken and

moved out

Track rib rear flange breaking

Bottom skin breaking following two spanwise directions

7. LE Repairs against bird impact 7. LE Repairs against bird impact –– Experimental testsExperimental tests



Num. Prediction Test Bird weight

Impact bay

Next bay Passed through

Total rest in the struct.

1.814 kg

0.274 kg

0.530 kg

Not measured

0.804 kg / 44%

1.814 kg

0.414 kg

0.650 kg

0.75 kg

? kg / ?%Test damage / cracked fasteners

- 2 principal cracks in the patch & top skin

- most of “intact” rivets

Num. prediction / cracked fasteners- 2 principle cracks

- 1 additional at patch limit

- most of intact rivets TR2 Sheetmetal rib

UpperGirder

Comparison between experimental and numerical results7. LE Repairs 7. LE Repairs –– ComparisonComparison



•• Parameters examined are shell thickness, curvature of Parameters examined are shell thickness, curvature of the LE profile, rivet pitch, patch size and backing plates the LE profile, rivet pitch, patch size and backing plates design. design. •• From the parametric studies occurs that the energy From the parametric studies occurs that the energy absorbing capability of the repair decreases with decrease absorbing capability of the repair decreases with decrease in thickness, decrease in curvature radius, dividing of in thickness, decrease in curvature radius, dividing of backing plates and increase in rivet pitch.backing plates and increase in rivet pitch.

0%

10%

20%

30%

40%

50%

60%

1 1.5 2 2.5 3Thickness [mm]

Ener

gy a

bsor

ptio

n [%

] RepairIntact

0%

10%

20%

30%

40%

50%

60%

70%

0 50 100 150 200 250 300

Radius [mm]En

ergy

abs

orpt

ion

[%]

RepairIntact

7. LE Repairs against bird impact 7. LE Repairs against bird impact –– Sensitivity analysisSensitivity analysis



8. Validation 8. Validation ExperimentalExperimental TestsTestsFFP1 : Fatigue flat riveted panel with skin repairsFFP1 : Fatigue flat riveted panel with skin repairs

FFP2 : Fatigue flat welded panel with welded stringer repairs anFFP2 : Fatigue flat welded panel with welded stringer repairs and dentsd dents

FCP1 : Fatigue curved riveted panel with skin repairsFCP1 : Fatigue curved riveted panel with skin repairs

FCP2 : Fatigue curved welded panel with welded stringer repairs FCP2 : Fatigue curved welded panel with welded stringer repairs and dentsand dents

SFP : Static flat welded panel with stringer repairsSFP : Static flat welded panel with stringer repairs

Objectives of the Panel’s Testing:

Q To validate the developed advanced calculation methods for fatigue life assessment and crack propagation of the investigated repair principles

Q To quantify the conservatism of pre-IARCAS calculation approaches for small skin repairs.

Q To support the justification and the extension of allowable damage limits



9. Conclusions9. Conclusions

• A significant reduction of operational and maintenance A significant reduction of operational and maintenance costscosts HAS been achieved through the improvement of HAS been achieved through the improvement of the Repair efficiency of aircraft metallic structuresthe Repair efficiency of aircraft metallic structures

•• FollowFollow--on investigations on the maintenance and on investigations on the maintenance and repair issues can contribute towards the goal of a more repair issues can contribute towards the goal of a more Affordable and Safer aircraftAffordable and Safer aircraft

recent developments for improve and assess repair ... · pdf filevienna, 19-21 june 2006...

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