Part1 Damage Tolerance Overview[1]

BYJ30-SDT-P00-00x -101/01/00/SAM

Structural Damage Tolerance — Part 1 —

Concepts and Overview

Matthew MillerStructural Damage Technology

(425) 266-5091

Course Number: 6CV40051

BYJ30-SDT-P00-00x Page 1-201/01/00/SAM

Damage Tolerance Overview



Concepts and Overview Outline

IntroductionKey elements of damage toleranceKey elements of residual strengthKey elements of crack growthKey elements of damage detection and maintenance planningDocumentation



Crack in Pressure Bulkhead



IntroductionResponsibilities of Structures Engineers

Static strengthVerify that the static strength of undamaged structure meets load

requirementsDurability (fatigue)

Verify that structural detail fatigue quality matches the required quality to meet the Design Service Objective

Damage toleranceVerify that an economically feasible inspection program can be

implemented to detect fatigue, corrosion, or accidental damage before the residual strength of the structure falls below the required fail-safe load level



IntroductionDefinitions of Damage Tolerance

Structure is damage tolerant if damage can be detected and repaired before residual strength falls below the regulatory load level

Damage tolerance means that the structure has been evaluated to ensure that should serious fatigue, corrosion or accidental damage occur within the operating life of the airplane, the remaining structure can withstand reasonable loads without the failure of excessive structural deformation until the damage is detected

Damage tolerance is the attribute of the structure than permits it to retain its required residual strength for a period of use after the structure has sustained specific levels of fatigue, corrosion, accidental or discrete source damage



SafetyDamage tolerance• Residual strength• Crack growth• Damage detection

EconomicsDurability• Fatigue• Corrosion prevention• Maintenance

Introduction Basic Concepts



Introduction Durability and Damage Tolerance Analysis Philosophy

XCritical Crack

Length(Allowable Damage)

Detectable Crack Length

Damage Tolerance

Durability(includes fatigue, corrosion and manufacturing quality)

Flights, N

N = Inspection Frequency

Cra

ck L

engt

h, L

Crack Initiation

Crack Propagation

N N N

MDSO X FRF< 5% of Fleet

Damage Detection

Period

N N

Crack growth analysis

Residual strength analysis

Inspection program



Introduction Damage Tolerance Technology Standards

D6-24958 “Damage Tolerance Methods and Allowables” (Books 3 & 3A)

Commercial Airplanes standards are based on 30 years of transport airplane development and service experience

Comprehensive methods and allowables are based on traditional aerospace industry fracture mechanics techniques

Audited and approved by the FAA as acceptable means to comply with FAR 25.571 requirements



Introduction Damage Tolerance Methods and Allowables (Book 3)

PurposeAid in the development of a maintenance program to inspect for potential

fatigue damage in a fleet of airplanes

Approach optionsToo simple

Loss of credibility Loss of accuracy

Too complex Excess resources and flow time required Excess probability of human error Loss of visibility and control by management

Just right Sufficiently accurate Easy to use



Safety requires diligent performance by all participants

Airlines

IntroductionAirplane Safety

Design, analysis and

manufacturing

Maintenance and

inspection

Regulatory agencies

• FAA• JAA• CAA• CAAC• AR



Structural Loads

Limit loadHighest load encountered by a fleet during its operational life

Ultimate loadsLimit load exceeded by a factor of safety, usually 50%

Operating loadsLoads normally encountered in day to day operations

Regulatory fail-safe loadsUsually limit load or equivalentDetailed in FAR 25.571 and JAR 25.571



Stru

ctur

al s

treng

th

Normal operating loads

Period of service

Structure is damage tolerant if damage can be detected and repaired before residual strength falls below the regulatory load level

Key Elements of Damage ToleranceDamage Tolerance Definition

Regulatory load requirement per FAR 25.571

Damage detection and restoration

Ultimate load capability required after damage detection and repair

NDI detection

period

Normal deterioration due to undetected damage

Detectable by NDI

Dam

age

size

Visual detection period

Visually detectable

Maximum allowable undetected damage



•Cyclic stress spectrum•Material•Geometry•Multiple Site Damage•Environment•Load distribution

Key Elements of Damage ToleranceDamage Tolerance Parameters

Crack growth resistance

Inspection program

Allowable damage

• Access / visibility• Inspection intervals• Inspection methods• Damage detection period• Multiple cracking in the

fleet• Damage detection

requirement

•Material•Geometry•Fracture toughness•Multiple Site Damage•Load distribution



+ + =

Key Elements of Damage ToleranceSummary

12

3

1. Residual strength analysis

2. Crack growth analysis

3. Inspection program

Safety of flight

Lcritical

Cra

ck le

ngth

Lcritical

Inspectable crack length vs. flights to

critical curve

Skin crack

Chord crack

Flights to critical

3

2

1Chord crack

TOTAL DTRBased on assumed inspection program

REQUIRED DTRBased on successful service experience

Compare Damage Tolerance Rating

Are additional inspections required?



Key Elements of Damage ToleranceDefinitions

Primary structureThat structure which carries flight, ground or pressure loads

Secondary structureThat structure which carries only air or inertial loads generated on or

within the secondary structureStructural Significant Item (SSI)

Any structural detail, element or assembly judged to be significant because its failure reduces airplane residual strength or results in the loss of function

Principal Structural Element (PSE)Structure that contributes significantly to the carrying of flight,

ground and pressurization loads, and whose failure could result in catastrophic failure of the airplane (Ref. AC 25.571 - 1b)



Key Elements of Damage Tolerance Structural Classification

All p rim ary structure notincluded in two

categories above

Structural CategoryTechnology

ControlMethod

Technique ofEnsuring Safety

S econdaryS tructure

OtherStructure

D esign for sa fesepara tion or loss o ffunction

C ontinuedsafe fligh t

D am ageO bvious orM a lfunction

Evident

Adequate residua lstrength w ith extensivedam age-obviousduring wa lkaround orind icated bym a lfunction

R es idualstrength

R es idualstrength

Crack growthInspectionprogram

Inspection programm atched to structu ra lcharacte ristics

D am ageDetection

by P lannedInspection

S afe-L ifeC onservative fa tiguelife Fatigue

StructuralClassification

Examples

W ing spoile r segm ent(sa fe separation or sa fe

loss o f function)

Typ ica l w ing sk in /stringersurface (fue l leak)

Landing gear s tructure

Dam

age

Tole

rant

Des

ign

Safe

-Life

Des

ign

Stru

ctur

ally

Sig

nific

ant I

tem

s (P

rimar

ySt

ruct

ure)



Obvious Damage!



Ground Accident



Key Elements of Damage Tolerance FAR/JAR 25.571

Certification requirements pre & post Amendment 45

Analysis Old FAR 25.571(Pre-1978)

New FAR 25.571(Post-1978)

Residualstrength

Single element orobvious partial failure

Multiple active cracks

Crack growth No analysis required Extensive analysis required

Inspectionprogram

Based on service history FAA air carrier approval

Related to structural damagecharacteristics and pastservice history

Initial FAA engineering and aircarrier approval



• Redundant load path.• Capable of

withstanding regulatory loads with a single element failed

(Pre - 1978) (Post -1978)

FAIL SAFE DAMAGE TOLERANT SAFE LIFE

Key Elements of Damage Tolerance FAR/JAR 25.571 Certification Requirements

• Capable of withstanding regulatory loads with partial or failed element with the presence of cracks in adjacent and attaching structure

• Able to detect and repair damage prior to structural strength loss below regulatory load capability (Amendment 45)

• Allowable crack size is very small (undetectable)

• Adequate inspections are impractical

• Parts are removed from service when “safe life” has been reached



Principal Damage SourcesPrincipal Structural Elements

Fatigue damage

Environmental damage

Accidental and discrete damage

Damage tolerance

Environmental damage evaluations

Approved inspection and maintenance program

Damage detection evaluations

Types of inspections

Typical accidental

Discrete source

Corrosion prevention plan

Corrosion inspection

Safe life

Accidental damage evaluations



1

1

Fatigue cracking is assumed to have occurred in the most difficult inspection/access location independent of calculated fatigue life or full scale test results



DETECTABLE

Key Elements of Damage Tolerance Inspection and Maintenance Philosophy

Environmental deterioration and accidental damage

Supplemental fatigue inspections

Flee

t dam

age

rate

Detectable size fatigue damage

Detectable fatigue

damage

Design Service ObjectiveYears of service

Thresholds are based on:• Fatigue approach for 727,737,747.

Typically 50% to 75% of DSO• Initial flaw approach for 737NG, 757,

767, 777 and any new program

Corrosion prevention and control program inspections

Scheduled maintenance check intervals

Fleet actions for WFD

Mandatory SB modifications

Repair assessments / inspections



+ + =

Key Elements of Damage Tolerance Residual Strength

12

3




Safety of flight

Strength? Flights? Inspections?



Static behavior

Transition behavior

Regulatory Requirement

LEFM behavior

Maximum allowable damage

Key Elements of Residual Strength Residual Strength Parameters

L

GeometryGeometry Correction

Factor

Y B

Thickness

MaterialFracture Toughness

KAKA

Y

Crack Length

Res

idua

l Stre

ngth



Stress,

L

K is the defining parameter for crack growth and residual strength predictions

Key Elements of Residual StrengthStress Intensity Factor

tk

CYnLK 3tk

Stress,

3

Stress concentration factor Stress intensity factor



Key Elements of Residual Strength Fracture Toughness

Fracture toughness is a function of:ThicknessOrientationAlloy, temper & formTemperature

0

10

20

30

40

50

60

70

80

0 0.2 0.4 0.6 0.8 1

Thickness (in)

KA (k

si

in)

Plane stress

Plane strain

Mixed mode (plane stress and plane strain)

KIC

7150-T77511

Extrusion

L-T orientation



Key Elements of Residual Strength Failure Criterion - Plane Stress Fracture

2L

ionsetoK:A

2L

cAoK:C

c

i

Lo Lc

Crack Length, L

A

BC

L

KA is the apparent fracture toughness associated with initial crack length at maximum load

KC is the fracture toughness associated with actual crack length at maximum load

Konset, KC and KA are not constants like KIC Most practical is KA - constant to first

approximation for given t

KA

Konset

KC

A

BC

2L

cCcK:B

Engineering approach to fracture assumes failure occurs when K = KA



Key Elements of Residual Strength Failure Criterion - Plane Strain Fracture

c

c

Lo LcCrack Length, L

Little or no plasticity involved when brittle failure occursFlat fracture surface

KIC

2L

cICcK

Failure occurs when K = KIC

L



Key Elements of Residual Strength Failure Criterion

Applies to all cracks in Titanium and all through thickness cracks in Aluminum and Steel

For a single corner crack (e.g., a lug) 0.05 inches or less, or two corner cracks (e.g., a fastener hole) 0.035 inches or less in Steel (< 240 ksi), use static strength analysis. For corner cracks of 0.01 Inches or less in Steel (> 240 ksi), use static strength analysis

For corner cracks in Aluminum 0.02 in. < L < thickness, use a straight line fit between the static strength at L= 0.02 inches and residual strength for L=thickness

Use typical yield stress; when data is not available use 2.228*”B basis” - 1.228*”A basis” or 1.10*”B basis” yield stress

AN/AG is the Net Area / Gross Area ratio for the damaged structure

(I/c)N / (I/c)G is the Net / Gross section modulus ratio for the damaged structure

In both the above cases, the Net properties = Gross - (holes & cracked out) properties

The “” is a shape factor for transition failures, = 0.63 for all materials

1

2

3

4



F FAA

FW L

WRS TYN

GTY

F

KL Y 1

, Y 1.99 0.41LW

...RSA

Key Elements of Residual Strength Failure Criterion - Transition Mode

Net section yield Linear Elastic Fracture Mechanics

B =0.5 inW =10 inMat’l 2024-T351 (L-T)KA = 125 ksi inFTY = 55 ksiW

B

L

Crack

0

10

20

30

40

50

60

70

80

0 2 4 6 8 10

Crack length (in)

Res

idua

l str

engt

h (k

si)

Net section yieldLEFM

Transition equation

Full static strength

L=t

Edge Cracked Panel Example



Cracking Patterns Recommended crack configurations Based on experience and engineering judgement



Y Factor Geometry correction factor in the stress intensity factor calculation Uses the superposition principle to include one or more geometry effects (J factors)



J and C FactorsJ factors account for differences between structure being analyzed and infinite plate

C factors account for increase in load due to cracking of attached structure



Y and C Factors

L L L

2LK YK 2

L CYK 2L

0.1Y

0.1C

0.1JY twidtheffec

0.1C

0.1Y

0.1C



+ + =

IntroductionKey Elements of Damage Tolerance

12

3




Safety of flight

Strength? Flights? Inspections?



Crack Growth ConceptsCrack Growth Parameters

X

Flights

Cra

ck le

ngth

Li (detectable)

Lf (critical)MATERIAL

GEOMETRYFLIGHT STRESS PROFILE

M

Y S

Stre

ss



1.E-07

1.E-06

1.E-05

1.E-04

1.E-03

1 10 100

Kmax (ksi in)da

/dN

(in/

cycl

e)

1

p

M

Crack Growth FundamentalsCrack Growth Equation

pmax4

MKZ10

dNdL

n1

Boeing Crack Growth Rate Equation (1979)

Based on P. C. Paris, 1961, and K. Walker, 1970, equations

M & p = Material crack growth rate parameters

Measures relative material resistance to crack growth

Reflects effect of environment

TEST

COUPON

0.1R00.1R1.1

0.0R0.1R1.010.1R0.0R1

Z

q



Allowable MSection 5, Book 3

Alloy Temper FormGrain

Orientation

Finished

Gage

(inches)

Allowable

M

Airplane Environment

W C B F

2014

2024

2219

2224

2324

T652T652

T3T4T3

T351

T3511T851T851

T6

T3511

T39

Hand ForgingDie Forging

Bare Sheet

Clad SheetPlate

Extrusion

Bare Sheet

PlatePlate

Die Forging

Extrusion

Plate

T-ST-L

L-T, T-LL-T, T-LL-T, T-LL-T, T-L

L-T, T-L

L-T, T-L

T-LL-T

L-T

L-T

.75

.75

.04 - .25

.04 - .06

.04 - .13

.10 - .15

.15 - .35

.35 - .50

.10 - .35

.37 - 1.0

.37 - 1.0

-

.16 - .31

.10 - .15

.15 - .35

.35 - .50

21.721.7

24.825.626.026.924.522.626.324.320.2

18.3

26.8

26.925.925.0

24.924.9

27.928.829.230.227.625.429.627.723.1

20.6

30.1

30.329.228.1

20.620.6

23.724.524.825.623.321.525.023.419.4

17.4

25.5

25.624.723.8

22.222.2

25.125.926.327.324.822.926.721.417.9

18.6

27.2

27.326.325.3

.

...

.

...

.

...

.

...

.

.

(cold)(wet) (bilge) (fuel)



Integration of Crack Growth Rate Equation

dLCYnL

n1

fZM10N

f

o

L

L

pp

max

4

pmax4

MKZ10

dNdL

n1

If loading is constant amplitude, the rate equation can be integrated and solved as follows for the number of cycles:

CYfK nL

maxmax

Note: Constant Z and fmax

X

N, number of cycles

Cra

ck le

ngth

Lo

Lf (critical)

fmaxS

tress

fmin

CdNMfZ10dLCY

n1 p

max4pn

L




Lower wing skin stress history for a “typical” flight

Stress(ksi)

Time

If loading is variable amplitude, and the crack growth rate does not depend on the sequence of varying load cycles, then the life integral can be discretized as shown:

ground

flight

C

o

N

0i

pimaxFp

4L

L

pn

L fZNM

10dLCYn1

p/1N

0i

pimax

C

fZS

dLCYn1

SM10N

L

L

pn

Lp

p4

F

o

NC = Number of cycles in a flightNF = Number of flights

Define stress rating, S, as:



Load Sequence Effects on Crack Growth

Overloads produce crack growth retardation, which is caused by the presence of residual compressive stresses in the yielded zone ahead of the crack tip

Source: Broek, D. , Elementary Engineering Fracture Mechanics, 4th ed., Kluwer Academic Publishers, Dordrecht, 1986

N, number of cycles

Cra

ck le

ngth




Lower wing skin stress history for a “typical” flight

Stress(ksi)

Time

To account for the effects of overload and underload, every stress cycle is adjusted, similar to the Willenborg model

ground

flight NC = Number of cycles in a flightNF = Number of flightsfmax eff = Effective fmax for cycle iZeff = Uses effective fmax and fmin for cycle i

Define spectrum stress rating, Sspectrum, as:

dLCgYn1

SM10N

L

L

pn

Lpspectrum

p4

F

o

p/1Ns

0i

pieffmaxeffspectrum fZS



Interfaced Durability and Damage Tolerance Analysis Software (IDTAS)

Complex flight operating stress profiles are analyzed using IDTAS

Loads and flight segment spectra are input, and relative fatigue damage and crack growth are calculated



IDTAS Damage Tolerance Output

SspectrumSspectrum/Slinear

IDTAS is the source of the stress rating, S, for use in the crack growth equation



Damage Detection and Structural Maintenance Planning

ObjectiveMaintain acceptable level of structural airworthiness throughout

the operational life of the airplane

GuidelineAirline / manufacturer maintenance program planning document,

MSG-3 (Maintenance Steering Group) Air Transport Association of America, Oct. 1980 (Rev. 2005.1)FAA approved as a means of complying with FAR 25.571




MSG-3 Rating Systems

Rating system provides quantitative means of determining inspection requirements

Development of rating system is responsibility of manufacturerRating systems are based on past practice and manufacturer /

operator experience with similar structureRatings address three principal sources of damage

FatigueCorrosionAccidental damage




Structural Maintenance Planning

Initial inspection program development and documentationThe initial structural maintenance plan for a new model is directed

toward detecting corrosion, stress corrosion and accidental damage

A feasibility study for fatigue damage detection is made to determine if a practical program can be constructed

Supplemental inspection programAs the fleet matures, the risk of fatigue cracking increases and the

inspection program is reassessed at 10 to 15 yearsIf the initial inspection program is inadequate for finding fatigue

cracks in significant structure, additional or supplemental inspections will be required




Environmental Deterioration Rating (EDR) System

Exposure to Adverse EnvironmentsSusceptibility Index

Probable Possible Unlikely

Standard 0 1 2

Proven /improved 1 2 3Environmental

ProtectionSpecial

attention 2 3 4

Sensitivity to Damage SizeTimely Detection Index

High Medium Low

Poor 0 1 2

Adequate 1 2 3

Visibility of theSSI for inspectionduring scheduled

maintenancechecks Good 2 3 4

EDR = Susceptibility Index + Timely Detection Index



Damage Detection and Structural Maintenance PlanningAccidental Damage Rating (ADR) System

Likelihood of Accidental DamageSusceptibility Index

Probable Possible Unlikely

Low 0 1 2

Medium 1 2 3

Estimated residual

strength after accidental damage High 2 3 4

Sensitivity to Damage GrowthTimely Detection Index

High Medium Low

Poor 0 1 2

Adequate 1 2 3

Visibility of theSSI for inspectionduring scheduled

maintenancechecks Good 2 3 4

ADR = Susceptibility Index + Timely Detection Index



Structural Maintenance PlanningKey Considerations and Responsibilities

for EDR and ADR EvaluationsApplication Primary ResponsibilityRating

Category EDR ADR Operator BoeingKey Considerations

Visibility Visibility for inspection after access

Sensitivity todamage size

or growth

Relative sensitivity within zoneconsidered

External: multiple element damage Internal: single element damage

Environmentalprotection

Comparison with previous protectionsystems and recent service history

Corrosion experience in same zoneExposure toadverse

environment

Material susceptibility to stress corrosionand potential for preload

Likelihood ofaccidentaldamage

Operator experience in the same zone

Strength afteraccidentaldamage

Likely size of damage relative to criticaldamage size



EXAMPLE

Structural Maintenance PlanningExample of EDR/ADR Application

Environmental deteriorationrating (EDR) or Accidental

Damage Rating (ADR)External Internal

1 2A 2A

2 5A 5A

3 3C C

4 2C

5 4C

6 or greater

2CAGE exploration: 4C interval

with 1/5 or 1/10 of fleet

Frequency of Structural Inspections ATA: Major Zone 300 (Section 48 & Empennage)

A-check frequency: 300 flight cyclesC-check frequency: 15 months or 3000 flight cycles, whichever comes first



PDProbability of

detecting damage

Probability of inspecting an aircraft with

damage

Probability of inspecting detail

consideredProbability of

crack detection

Structural Maintenance PlanningDamage Detection Parameters

P1 P3P2



Structural Maintenance PlanningDefinition of Inspection methods

InspectionMethod

Description for Boeing DTR System(See MSG-3)

TypicalApplications

Walkaround Observations from the ground to detect obviousdiscrepancies such as fuel leaks (see category 2 structure) Pre-flight

General visual(no MSG-3 credit)

Visual check of exposed area of wing lower surface, lowerfuselage, doors and door cutouts, and landing gear bays A-check

Surveillance(MSG-3 general

visual equivalent)

Visual examination of defined internal or external structuralarea from a distance considered necessary to carry out anadequate check. External includes structure visiblethrough quick-opening access panels or doors. Internalapplies to obscured structure requiring removal of filletsfairings, access panels or doors, etc. for visibility

C-check

Detailed(MSG-3 detailed

equivalent)

Close intensive visual inspections of highly definedstructural details or locations searching for evidence ofstructural irregularity

D-checkselected items

Special (seeBook 3 pg. 12-26)

Inspections of specific locations or hidden details usingspecified nondestructive inspection (NDI) procedures.

D-checkselected items



Structural Maintenance PlanningDamage Detection Period

Crack Growth Interval, Flight Cycles

Cra

ck L

engt

h

Damage Detection Period, Flights to Critical

Cra

ck L

engt

h

CriticalCritical

Surveillance

Detailed

NDI

Surveillance

Detailed

NDI

Engineering Format Maintenance Planning Format



Maintenance PlanningDetermining Probability of Detection Parameters for Inspections

Flights

N N NIn

spec

tabl

e C

rack

Len

gth

Need to relateInspection methodProbability of detectionCrack length

Safe damage detection period

0.0001

0.001

0.01

0.1

0.99

0.1 1 10

Inspectable Crack Length (in)

Pro

babi

lity

of D

etec

tion

Detailed

Surveillance

Audited by the FAA in 1980 Inspectable Crack Growth Curve

General Visual

n

1iDD i

P̂11P

m

1j

n

1iDD ij

P̂11P

For n inspections of the first crack

For n inspections on m cracked airplanes



Relative Damage Detection Reliability for Visual Inspections

0.9

0.95

0.99

5 10

Inspectable Crack Length (in)P

roba

bilit

y of

Det

ectio

n

303 7

0.0001

0.001

0.01

0.1

1

0.1 1 10

Inspectable Crack Length (in)

Pro

babi

lity

of D

etec

tion

3050.5



Inspectable Crack Length

Inspection access: Top

Inspection access: Bottom

A

CB

Cra

ck L

engt

h Critical

Actual Crack Growth Curve

FlightsNA NBNC

AC

B

X

Insp

ecta

ble

Cra

ck L

engt

h

Inspectable Crack Growth Curves

FlightsNA NBNC



When do Supplemental Inspections Start?

Current inspection threshold determinationBased on crack growth from an equivalent manufacturing flawFAA recommended (used for 757 and 767)

Initial flaw size regardless of material or manufacturing process of– 0.05” for open holes– 0.03” for filled holes– 0.005” for cold worked holes

Threshold life is crack growth life from initial flaw to critical divided by 2FAA mandated 100% inspection after reaching thresholdBoeing new method (used for 777,

737NG and new models and derivatives) Establish initial flaw sizes based on

fatigue quality of detail Crack growth life to detectable or

through thickness

DFR

Equivalent Manufacturing Quality Flaw Size



Documentation for a New Airplane Program

Structural Inspection Planning Data Document (SIPD) Description of structure Completed EDR/ADR forms DTR forms from feasibility study

Maintenance Planning Data Document (MPD)Maintenance Review Board (MRB) Report



Reported Discrepancies

Discrepancies to be reported All cracks and previously unreported occurrences of significant corrosion involving any

SSI or PSE shall be reported promptly to Boeing Boeing follow-up actions

All adverse SSI or PSE reports will be reviewed immediately to determine if they are findings directly applicable to the SSI being addressed.

If discrepancies are related to fatigue cracking of the SSI, the following actions are to be taken

– Report factual data available on finding, with appropriate priority, to operators and airworthiness authorities

– If required, develop a fleet inspection program to obtain additional data necessary to formulate a service bulletin

– Prepare and issue a service bulletin addressing recommended structural modifications and inspections

– Revise the Supplemental Inspection Program Document (SIPD) to eliminate this SSI or a portion of the SSI, and reference the new service bulletin in the safety of flight service bulletin list



Supplemental Structural Inspection ProgramsMethods are based on current practice, accessibility, critical crack length and non-destructive inspection

procedures

707/720 SSIP 727/737/747 SSIP

DamageTolerance

Assessment

Assessment of all SSI's(with and without a historyof in-service fatiguecracking) assumingcracking may have occurred

Assessment of all SSI's (withno history of in-servicefatigue cracking) requiringinspection to ensure timelydetection of fatigue damage

InspectionProgram

Intervals determined bydividing life (from detectableto critical) by a scatter factor

Programs augment or modifyexisting maintenance plansand intervals are based onlife (from detectable tocritical), multiple inspectionsand multiple fleet cracking)

AirplaneEffectivity

All airplanes exceeding thethreshold for an SSI aresubject to the requirementsfor that SSIDetails with known fatigueproblems have post-terminating inspectionrequirements with uniquethresholds based on servicedata

Only candidate fleetairplanes subject to in theinspection programrequirementsDetails with known fatigueproblems covered by existingservice bulletins arereassessed with inspectionsadded as needed



Need a quantitative measure of the probability of detecting a crack

Calculated DTR vs. Required DTR

DTR = Damage Tolerance Rating

Number of 50/50 chances of detecting damage

Maintenance PlanningDamage Detection Rating System

Structure Required DTR

Externally visible areas 4

Wings and nacelles Areas not externally visible 6

Primary flap structure 8

Empennage Primary structure 6

FuselageContribution of

cabin differential<50% 6

pressure to total fail-safe stress >= 50% 10

Part1 Damage Tolerance Overview[1]

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