F/A-18E/F FULL SCALE STRUCTURAL FATIGUE TESTING

F/AF/A--18E/F FULL SCALE STRUCTURAL 18E/F FULL SCALE STRUCTURAL FATIGUE TESTINGFATIGUE TESTING

Timothy N. CallihanThe Boeing Company – Integrated Defense Systems

Aircraft Structural Aircraft Program – December 6, 2007Palm Springs, CA

Timothy N. CallihanThe Boeing Company – Integrated Defense Systems

Aircraft Structural Aircraft Program – December 6, 2007Palm Springs, CA

Approved for Public Release, 265XPR-059.05 2December 6, 2007

The Boeing Company

FullFull--Scale Fatigue TestingScale Fatigue Testing• Test Articles

– FT50 represents the entire F/A-18E/F airframe– FT76 represents the new F/A-18E/F forward fuselage– FT77 represents the new F/A-18E/F wing

• Test Requirements– FT50 & FT77: Complete 18,000 simulated hours of fatigue cycling (3 lifetimes)– FT76: Complete 12,000 simulated hours of fatigue cycling (2 lifetimes)

Approved for Public Release, 265XPR-059.05 3December 6, 2007

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TensionTension--Compression Load SystemCompression Load System

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FT50 Fatigue Test SetupFT50 Fatigue Test Setup

14.1Maximum Lines Per Minute

1560Continuously Monitored Data Channels

27Pretest Surveys

5,786,042 (including make-up cycling)Total Spectrum Lines in 2 Lifetimes

448,492Lines per 1000 Simulated Flight Hours

182Load Controllers

89Deflection Transducers at Test Start

1643Strain Gages at Test Start

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FT50 Fatigue Test SpectrumFT50 Fatigue Test Spectrum• Symmetric maneuvers included both steady state and abrupt

conditions from -2.5g to 8.5g

• Abrupt maneuvers only occur for positive Nz events

• Asymmetric maneuvers included 40%, 60%, 80%, and 100% lateral stick deflections and consist of -1g rolls, +1g rolls, and Rolling Pull-Outs (RPOs) up to 8.5g

• Negative 1g rolls were limited to rolling through 180 degrees and positive 1g rolls continued through 360 degrees

• Rolling angles (bank to bank) were dependent upon the entry g level for RPOs at Nz greater than 1g

Approved for Public Release, 265XPR-059.05 6December 6, 2007

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FT50 Fatigue Test SpectrumFT50 Fatigue Test Spectrum• Ground-Air-Ground Cycles 15750

• Field Taxi Runs 8750

• Catapult Launches 2250

• Landings– Arrested 2250– Touch and Go 450– Field Carrier Landing Practice (FCLP) 6450– Field Mirrored Landing Practice (FMLP) 6600

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Extensive Assessment Of LoadingExtensive Assessment Of LoadingAccuracy of FullAccuracy of Full--Scale Fatigue TestScale Fatigue Test

24,000

Design

N Test

24,000LR = Life Ratio =

CI Life (SFH)

KtX

Sigm

a

Test 2(undertest)

Test 1(overtest)

N Test1 N Test2

Overtest if LR > 1.15Undertest if LR < 0.85

N Test

12,000TM = Test Margin =

Design Life:FT50 = 16,000FT77 = 16,000FT76 = 24,000

Test life at the design stress level

Approved for Public Release, 265XPR-059.05 8December 6, 2007

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Results of FT76 Fatigue Test AccuracyResults of FT76 Fatigue Test Accuracy

Y204.5

Z111.65

Z108.65

Y383.

Y260.7

Y326.5

Y357.

Y286.

Y233. Original FT50 DataLife RatiosOvertest: LR > 1.15EquivalentUndertest: LR < 0.85Test MarginsElevated Risk: TM < 1.33Minimal Risk: TM >=1.33

LR 0.69TM 7.65

UPR007E

LR 0.86TM 21.05

UPR009E

LR 0.80TM 2.38

UPR010E

LR 1.09TM 31.65

UPR021E

LR 1.60TM 2.24

UPR022E

LR 1.41TM 22.54

UPR026E

LR 1.33TM 1.77

UPR027E

LR 0.71TM 3.89

UPR038ELR 0.71TM 13.20

UPR039E

LR 1.50TM 13.73

UPR043ELR 0.62TM 5.79

UPR048ELR 0.08TM 72.10

UPR051E

LR 0.35TM 4.38

UPR052ELR 0.62TM 10.98

UPR080E

LR 0.77TM 11.51

UPR087E LR 0.62TM 4.00

UPR159E

Approved for Public Release, 265XPR-059.05 9December 6, 2007

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Remove Test Article fromFixture

Remove Wings, ControlSurfaces, and Vertical Tails

OML Inspection

Remove Doors andCovers

Available for visualinspection on request

Major Disassembly

NDI

Failure Analysis

Storage

Part Removal

FT50 Teardown Overview

Orientation and Training

Document visual anomalies

Document comparison tobaseline inspection results

Eddy Current used as the primary NDI technique

Looking for cracks that are 0.030” or larger

Extensive coordination with United States Navy to define expectations and processes

Approved for Public Release, 265XPR-059.05 10December 6, 2007

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FT50 keel longeron Y555 cracks FT50 keel longeron Y555 cracks at 10,000 SFHat 10,000 SFH

Approved for Public Release, 265XPR-059.05 11December 6, 2007

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FT50 rightFT50 right--hand outer wing electrical hand outer wing electrical cutcut--out cracks at 16,410 SFHout cracks at 16,410 SFH

Approved for Public Release, 265XPR-059.05 12December 6, 2007

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FT50 Leading Edge Extension (LEX) FT50 Leading Edge Extension (LEX) cracks at 15,000 SFHcracks at 15,000 SFH

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1717”” Crack on Crack on FT50 Right Wing FT50 Right Wing

Root Lower Root Lower SurfaceSurface

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Failure analysis example (wing root failure)Failure analysis example (wing root failure)

Approved for Public Release, 265XPR-059.05 15December 6, 2007

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Detailed models were constructed to characterize Detailed models were constructed to characterize the stress field at the countersink holesthe stress field at the countersink holes

Approved for Public Release, 265XPR-059.05 16December 6, 2007

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Failure surface of crack emanating from flaw at Failure surface of crack emanating from flaw at fastener hole in wing root lower skinfastener hole in wing root lower skin

Approved for Public Release, 265XPR-059.05 17December 6, 2007

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• Probability analysis was performed early in the F/A-18E/F program to determine the scatter factor that was to be used during fatigue analysis of structural components

- A scatter factor of 1.0, i.e., design for 12,000 SFH and test for 12,000 SFH, would correspond to a 50% probability that a crack will initiate after two lifetimes (12,000 SFH) of testing

• The E/F program incorporated a scatter factor of 1.33 (design for 16,000 SFH) into all fatigue design allowables which corresponds to an 18% probability that a crack will initiate after two lifetimes

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• Stretch testing the FT50 test article to three lifetimes (18,000SFH) corresponds to an effective scatter factor of 0.89 since the structure was only designed for 16,000 but was tested for 18,000 SFH

• This scatter factor corresponds to a 65% probability that a crack will initiate after two lifetimes

• Therefore, the number of cracks identified on the FT50 test article is not unexpected

Approved for Public Release, 265XPR-059.05 19December 6, 2007

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Summary of cracks on FT50Summary of cracks on FT50at completion of 18,000 SFH of testing at completion of 18,000 SFH of testing

FT50 Component Parts TotalWing 43 641

Forward Fuselage 53 492Center Fuselage 133 1119

Aft Fuselage 56 293Mechanisms 2 3Vertical Tail 9 45

Totals --> 296 2593

Crack Summary

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Many Parameters Required To Many Parameters Required To Successfully Complete Test ProgramSuccessfully Complete Test Program

Must perform many tasks to achieve a successful test

CertificationCertification

Perform Over/Undertest Analysis

Failure Analysis CapabilityNDI Capability

Analytic Tools

Define Teardown Criteria

Coordinated Teardown PlanCustomer CoordinationStrain SurveysSt

ruct

ural

Tech

nolo

gy

SUCCESS

Approved for Public Release, 265XPR-059.05 21December 6, 2007

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F/AF/A--18E/F Test Conclusions18E/F Test Conclusions• The F/A-18E/F full-scale fatigue test programs are a solid

success as expressed by United States Navy– “Superb” technical performance expressed by customer– FT50 2nd lifetime testing completed ahead of schedule

• Comprised multiple full scale fatigue test articles that certified the airframe to meet the intended design life

• Successful in identifying structure that is capable of exceeding the intended design life

• Success of the F/A-18E/F airframe continues to be realized by utilizing the E/F structural platform as the foundation of the EA-18G program – Next generation of electronic attack aircraft for the United States Navy

Approved for Public Release, 265XPR-059.05 22December 6, 2007

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EAEA--18G18GElectronic attack variant of the F/AElectronic attack variant of the F/A--18E/F18E/F