f15 service life extension 2025-asip2010.pdf

WR-ALC/GRM-Eagle Division

December 02, 2010 | 1

Certifying the F-15C Beyond 2025Paul A. ReidThe Boeing Company

Joseph D. LaneWarner Robins-ALC/GRMEB

Aircraft Structural Integrity Program (ASIP) 2010 Conference

2

Outline

• Introduction

Evolution of F-15C service life

Service life projections

• Full-scale fatigue test history

• Current FTA7/8 full-scale fatigue test

Testing approach

Configuration

Engineering activities

FEM overview

Test spectrum overview

Test severity tool

Health monitoring

• Current FTA7 status


3

Air Combat Command plans to fly the F-15C beyond 2025Full-Scale Fatigue Testing Needed

Production run 1978 – 1986 Avg A/C will go thru PDM 7 times by 2025

Evolution of F-15C Service Life Contractual

Service Life

'76 - '80 FTA1

'81 - '93 FTA1

1994 Fleet Usage

'94 - '08 FTA6

'08 - '13 FTA6

'14 - '25 FTA7

Spectrum Flight Hours

8,000

1 Fail-Safe (Crack Initiation) and Safety Factor of 4.0

4,000 1

0.33

1.0

1.0

4,000 1

Spectrum

18,000 TBD 1.37

8,000 2

Economic Service Life is 16,000

18,000 9,067

0.72

9,067

3 Relative to Critical Hole on FS 626 Bulkhead Lower Cap

2 Damage Tolerance and Safety Factor of 2.0 Introduced

Time

Period

Spectrum

Severity 3

0.33

USAF Required

Service Life

Certified Service

Life

4,000 1 4,000


4

Required Service Life Projected to 2025Individual Aircraft Equivalent Flight Hour (EFH) Projection

21

Life PredictedDI CumYears#Year

DISquadron EFH Projected

Factors to consider:

• Squadron re-distribution

• Damage rates

• Future changes to location(s) of interest initial flaw size

continuing damage

preventive repairs

• Fleet size

1 DI = Damage Index, the percent of crack growth life depleted using individual aircraft usage

2 Crack growth life at location of interest, such as the critical economic life location

Describes individual aircraft usage in terms of a reference test spectrum


5

0

2,000

4,000

6,000

8,000

10,000

12,000

14,000

0 2,000 4,000 6,000 8,000 10,000 12,000 14,000

Reported Flight Hours

Eq

uiv

ale

nt

Fli

gh

t H

ou

rs

Current Equivalent Flight Hour SummaryAs of 31 December 2009

Assumptions:

2009 Squadron distribution

2009 Damage rates

Revised 626 bulkhead centerline analysis

FTA6 EFH

All aircraft

Aircraft exceeding 9,067 Hour

Certified Service Life*

* These aircraft have “restricted” airworthiness certificate

Life PredictedDI Cum EFH


6

2025 Projected Equivalent Flight Hours (EFH)2007 vs 2009 Estimate

F-15C/D Golden Fleet 2025 Projected Equivalent Fatigue Spectrum Hours (EFSH)

FS 626 Bulkhead at Centerline Location (Location 5000)

Average EFSH = 10847 hrs

0

2,000

4,000

6,000

8,000

10,000

12,000

14,000

16,000

18,000

20,000

0 2,000 4,000 6,000 8,000 10,000 12,000 14,000

Projected Actual Flight Hours

Pro

jec

ted

Eq

uiv

ale

nt

Fli

gh

t H

ou

rs

2007 Estimate

2009 Estimate

2025 Projected Equivalent Flight Hours (EFH)2007 Estimate

F-15C/D Golden Fleet 2025 Projected Equivalent Fatigue Spectrum Hours (EFSH)

FS 626 Bulkhead at Centerline Location (Location 5000)

Average EFSH = 10847 hrs

0

2,000

4,000

6,000

8,000

10,000

12,000

14,000

16,000

18,000

20,000

0 2,000 4,000 6,000 8,000 10,000 12,000 14,000

Projected Actual Flight Hours

Pro

jec

ted

Eq

uiv

ale

nt

Fli

gh

t H

ou

rs

2007 Estimate

2007 Assumptions:


2007 Damage rates

Historical FS 626 bulkhead analysis

FTA6 EFH

Long term fleet

2009 Assumptions:


2009 Damage rates

Revised 626 bulkhead analysis

FTA6 EFH

Long term fleet

2007 Assumptions:


2007 Damage rates

Historical FS 626 bulkhead analysis

FTA6 EFH

Long term fleet


7

F-15C/D Inventory Projections

0

50

100

150

200

250

300

350

400

450

500

1978

1988

1998

2008

2018

2028

2038

Year

Nu

mb

er

of

Air

cra

ft

FSFT CSL Goal of CSL=18000

C/D Inventory Needed by USAF

C/D Inventory Based on CSL=9067

C/D Inventory Flying Beyond CSL

Anticipated Test CSL


8

Full-Scale Fatigue Test HistoryFTA-1, FTA-3, FTA-5: A/B Model (Block 1)

1974 – 1976

Component Tests:

- Fwd Fuselage

- Rudder

- Speedbrake

- Landing Gear

- Flap/Aileron (FTA-5)

Test Result:

16,000 Hours

No major failures

Certified Safety Limit = 4,000 Hours


9

A/B Center

Fuselage

C/D Wing

A/B Wing

Aft Fuselage

(dummy)

Full-Scale Fatigue Test HistoryFTA-6: A/B/C/D Models

1988 - 1994 Test Result:

18,133 Hours

Major failure at intermediate spar

Certified Service Life = 9,067 Hours

Fatigue Test Teardown Resultsnumber of cracks found

FTA-1, 3, 5 FTA-6

Wing A/B 111 249

Wing C/D n/a 138

Forward Fuse. 0 n/a

Center Fuse. 1 126

Aft Fuse 17 n/a

Components 0 n/a


10

FTA-7/8 Full-Scale Fatigue Test2009 2010 2011 2012 2013 2014 2015 2016 2017 2018 2019

Engineering / Lab Prep.

Test Article Arrives

FTA7 Testing

FTA8 Testing

Phase II Test Article Teardown

Phase III Correlation Analysis

Phase I

2.1 Years

2.75 Years8 Months

4.66 Years3.9 Years

Mar

FTA-7

• Fwd. Fuselage

• Center Fuselage

• Aft Fuselage

• Wings (2 sets)

FTA-8

• Horizontal Stabilator

Test Article Configuration

Only Phase 1 funded

Loading FixtureFTA-7

FTA-7 FTA-7

FTA-8

Loading Fixture

Loading

Fixture

Loading Fixture

FTA-7

Test Article

• 30 years old

• Latest PDM March 2010

• 8,426 FTA-6 EFH (8,100 AFH)


11

FTA-7/8 Test Article ConfigurationPreventive Repairs

Installed on all aircraft during PDM

• Areas redesigned from cracks found in earlier tests

• New inboard spars installed

• Significant repair budget in place to maximize endurance

• Preventive Repairs Installed:

Upper outboard longeron at FS 502

Vertical stabilizer picture frame with Gridlock

Various simple fixes to holes that cracked in FTA6

Canopy sill longeron at FS 377

Doubler repairs on lower wing skin

FS 626.9 bulkhead lower cap


12

Similar view,Similar view,

Figure B-1

View looking Outboard

Zone 1 Zone 2

Forward

0.6875”


Zone 1 Zone 2

Forward

0.6875”


Zone 1 Zone 2

Forward

0.6875”


Zone 1 Zone 2

Forward

0.6875”

FS 626.9 Bulkhead Lower Cap

Fleet Failure

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0 2,000 4,000 6,000 8,000 10,000 12,000

FTA6 Spectrum Hours

Cra

ck

Le

ng

th (

in.)

Analytical Match of Fleet Failure Achieved

Hole 12B cracked in FTA6

from pit on forward side of hole

Forward

Fleet failure occurred in 11,543 FTA6 EFH

Forward

Outboard


13

View Looking Forward

FS 626.9 Bulkhead Lower Cap Preventive Repair

Bulkhead expected to crack:

• Cracks found in FTA6

• Cracks found in service

• Repair Concept: Reinforce with fittings


Permanent Repair


Temporary Repair

View Looking Forward

Additional FS 626.9 Bulkhead Repair Configurations


14

Project Engineering Support ActivitiesStability & Control• CFD airload distributions

Mass Properties• Section mass dist.

• A/C mass dist. For FEMLoads • PITS studies

• Load balances

• Aeroload pressure dist.

• Theoretical external loads

• 38 balanced test loads

• Actuator loadsStrength

• FEM development• Inertia load points

• Airload points

• FEM refinement

• Theoretical & test FEM loads

• Test article requirements

• Strain gage defn. / correlation

• Strain surveys

Fatigue• Test plan

• Strain gage definition

• Test severity tool

• Over/under studies• Master events spectrum criteria

• 38 test load case definitions

• Target spectrum

• Test spectrum• Spectrum generation software

Merc/Boeing FEM

Lab


15

Mercer Engineering Research Center (MERC) / Boeing Global FEM

• Full FEM, 914 K shell elements, ~6 million DOF

• Each part modeled with shell elements at mid thickness, average element size ~1”

• Parts connected with ~250 K rigid body elements (RBEs)

• Approximately 7000 individual parts modeled

MERC FEM, element boundaries turned off

Substructure detail


16

Boeing Efforts to Prepare FEM in Support of FTA7

Define inertial mass points:Estimated at ~ 6,000 nodes

Define discrete mass points:All non-structural items over 10 lbs, ~ 113 items

Create coarse aero shell

for CFD pressure mapping:Original mesh aero shell was > 300,000 elements

Boeing aero shell, ~ 30,000 elements

Tank 1

Create fuel cell shells

for fuel pressure mapping


17

Preliminary FEM Validation Plan:

• ~ 1,000 strain gage locations mapped to primary load paths

• Pre-strain survey check-out of all load cases

• Compare predicted vs measured test strains on select number

of gages at 80% load level.

• Correlation criteria is +/- 3 ksi for stresses < 20 ksi and +/- 15%

for stresses > 20 ksi

Boeing Efforts to Prepare FEM in Support of FTA7

Status of Model Changes Made to Date:

• Aero and fuel pressure shells created

• Addition of canopy and windscreen

• Addition of vertical tip pods

• Addition of rudders

• Changes to wing dihedral

• Replacement of honeycomb core with

Gridlock in wingtips, control surfaces, and vertical tail

• Replacement of horizontal tails with existing fine

grid mesh model

• Re-modeling of wing-to-fuselage joints

• Validation of properties

• Various structural idealization assumptions

• 38 balanced test and theoretical load cases

• Actuator unit load cases


18

FTA7 Spectrum Generation

• Nz usage based on last 5 years, adjusted to reflect future usage

expectations.

Effect accounted for using base “weighting” factors

• Usage parameters that define target spectrum:

Nz data from Counting Accelerometer (C/A)

Damage rates at key tracked locations

• 22 Symmetric Points-In-The-Sky (PITS)

• 16 Asymmetric PITS


19

1

10

100

1,000

10,000

100,000

-0.5 -0.3 -0.1 0.1 0.3 0.5 0.7 0.9 1.1

Spectrum Load Level

Ex

ce

ed

an

ce

s p

er

1,0

00

Flig

ht

Ho

urs

Peak-Valley Exceedance Comparison - Wing Root Bending

Validation of applied spectrum accuracy:

• FEM validation (strain surveys)

• Near real time external load controls severity tool

• Comprehensive over/under test severity tool

• Crack growth correlation, spectrum truncation, and marker band test program

Need to quantify how accurate test spectrum is being applied….

• Theoretical spectrum: Match base-weighted

targets using actual Structural Data Recorder

(SDR) files

• SDR files chosen using optimization routine

C/A Nz counts

Damage rates at tracked locations

Typical target error ≈ 2.5%

• Actual test spectrum: Match theoretical

spectrum with 38 balanced load conditions

FTA7 Spectrum Summary

Spectrum Severity

FTA1 = 0.33

FTA6 = 1.0

FTA7 = 1.37


20

1

10

100

1,000

10,000

100,000

-0.5 -0.3 -0.1 0.1 0.3 0.5 0.7 0.9 1.1

Test Severity Tool OverviewTheoretical Loads

Cond. 1, 2, …n

P1 P2 … P107 P108

Theoretical Target Actuator Loads:

P1T = 5000 lbs,

P2T = 1250 lbs,

…

P108T = -400 lbs

Actual Measured Actuator Loads:

P1A = 5075 lbs,

P2A = 1235 lbs,

…

P108A = -404 lbs

Strain Correlated FEM

Over UnderTheoretical

Build Up Theoretical and Actual Stress Spectra

P1 P2 P3 Pn

Cond 1 C11P1 C12P2 C13P3 C1nPn



Cond n Cn1P1 Cn2P2 Cn3P3 CnnPn

Determine Stress Influence Coefficients

Extract FBD / FEA Stress and

Perform Crack Growth Analyses

• External Load Controls

• Any Area of InterestCra

ck L

en

gth

Spectrum Flight Hours


21

The Delta Damage Ratio Since “X” Day(s) Ago is the

change in damage in “X” days (4/22/04 to 4/23/04 in this

example).

There are four options used to identify the window of interest:

1) end date

2) # of days before

3) cycling or calendar days?

4) the days’ first or last data file?Find actual chart and change comments

Near Real-Time External Load Controls

Severity Tool

- -TEST vs PREDICTED DAMAGE RATIOS

1.001.001.00

1.08

1.03

1.001.02

0.980.960.97 0.981.00

1.08

1.031.021.02

0.99

0.95

0.99

1.07

0.80

0.90

1.00

1.10

1.20

1.30

1.40

1.50

WRBM WRTQ WFBM WFTQ ILEFHM OLEFHM TEFHM AILHM WTFBM WTABM

EXTERNAL LOAD CONTROLS

DA

MA

GE

RA

TIO

Delta Damage Ratio Since 1 Day Ago from

(16,781 SFH to 16,952 SFH)

Total Cumulative Damage Ratio to Date from

(12,000 SFH to 16,952 SFH)

WRBM

P1 - P55

P73

P77 - P97

WRTQ

P1 - P55

P73 - P97

WFBM

P13 - P19

P27 - P39

P53

P55

WFTQ

P13 - P19

P27 - P39

P53

P55

ILEFHM

P23

P25

OLEFHM

P27

P29

TEFHM

P41 - P51

AILHM

P31 - P37

WTFBM

P53

P55

WTABM

P53

P55

FT77 CUMULATIVE TEST SEVERITY

THROUGH THE END OF EACH BLOCK (Control Points)

0.80

0.85

0.90

0.95

1.00

1.05

1.10

1.15

1.20

WRBM WRTQ WFBM WFTQ ILEFHM OLEFHM TEFHM AILHM WTFBM WTABM

CONTROL POINTS

DA

MA

GE

Block 13

Block 14

Block 15

Block 16

Block 17

Block 18

CYCLE RATES

Block 10: 14.0 lpm

Block 11: 13.2 lpm

Block 12: 14.2 lpm

Block 13: 16.7 lpm

Block 14: 18.3 lpm

Block 15: 18.5 lpm

Block 16: 18.5 lpm

Block 17: 18.5 lpm

Block 18: 18.5 lpm

Sample Output


22

Structural Health Monitoring SystemsPiezoelectric Transducers

1) Piezoelectric Transducers Drive and

Sense Acoustic/Ultrasonic Energy

Transmission Through Structure

12

43

2) Propagation Paths are

Disturbed by Structural Damage

Causing Changes in Propagation

Damage Depth Prediction from Scatter Algorithm

SHM Estimated Damage Depth = 0.187”

Scattered Image Volume

Impact 15

Impact 15, Flat Panel 42 Ply 3) Multiple Algorithms provide

highly sensitive damage

indicators up to and including

3D damage imaging

Installed on:

• Intermediate Spar Lower Cap

• Shoulder Rib

• FS 626 Bulkhead Preventative Repair


23

Structural Health Monitoring SystemsComparative Vacuum Monitoring (CVMTM)

Strip Sensor

Bolt HoleSensor

1) A sensor has a matrix of separated alternating

galleries a Vacuum (red) gallery and an Ambient (blue)

gallery which are open to the surface to which they are

adhered to.

Installed on:

• Intermediate Spar Lower Cap

• Shoulder Rib


24

FTA7 Fatigue Test Status:October, 2010

Test Loads Development:

• Mass properties 100%

• CFD solutions 100%

• Balanced loads 30%

• Actuator loads process 65%

• Actuator balanced loads 0%

Spectrum Development:

• Master events criteria 100%

• Usage spectrum 100%

• Test load conditions 100%

• Preliminary test spectrum 25%

• Spectrum generation software 75%

• Spectrum validation testing 20%

• Test severity tool 10%

Finite Element Model:

• FEM refinement 67%

• Strain gage prediction/correlation 0%

Lab:

• Strain gage drawings 100%

• Instrumentation installation 50%

• Wing loading pad install 50%

• Fuselage load fitting install 60%

• Test fixture 25%

Test Cycling Start Date:

Sept. 07, 2011

f15 service life extension 2025-asip2010.pdf

Documents

external load

delta damage

cracks found

boeing efforts

prepare fem

safety factor

shell elements

theoretical