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UNCLASIFIED

The DASA & AD (AS) ASI Journey to date

L. Molent AM

ADFASI Symposium Mar 2019

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In a RAAF Officers Mess in a near by Universe…

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AcknowledgementsMake a Difference: Adopt a DST Scientist!

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Overview

DST AD & Aircraft Structures Branch (ASB) have a long history of partnering with DASA

We have always tried to insert innovation whilst maintaining schedule for delivery of ASI results

Lead to many innovative tests and analytical tools

AD tries to bridge the scale. i.e. from material coupon data to full-scale; importantly also from imperial to science (e.g. atomistic modelling (i.e. many lifing tools are imperial in nature).

To be successful AD needs a regulator on-board

– Industry & partners are important collaborators

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ASI: More than FatigueAircraft

Structural

Integrity

Aircraft Structures

Aeroelasticity &

Dynamics

Aero Loads

Exp Stress Analysis & Sensors

ASE

APS

Probabilistic Methods

AMTForensic Investigations

MD

Material Data

Analytical crack growth modelling

Composite Behaviour & Repair

Structural Analyses

Non-Destructive Inspection Research

Paints & Sealants

Environmental Protection

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AD: From Atoms toAircraft

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3

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67

8

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10

8b

𝑑𝑎

𝑑𝑁=

𝐷(∆𝐾 − ∆𝐾𝒕𝒉𝒓)

𝟏 −𝑲𝒎𝒂𝒙𝑨

𝟐

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RAAF PC-9/A PEARLA & FALKOR

PC-9/A Fin

Pc-9/a Empennage and Aft-fuselage Recertification and Life Assessment (PEARLA) launched by RAAF late 2010 to address fleet issues:

→ RAAF PC-9/A fleet condition data showed that the a/c experienced unexpected or premature fatigue cracking in aft fuselage primary structure

→ No operational load measurement (OLM) program conducted on aft fuselage

→ KNOWN aft fuselage FSFT load spectrum deficiencies

Proposed Approaches

→ Enhanced teardown of fleet aircraft: No airframe avail

→ Certify analytically viz FEM like FAR23 (Help regulator!!)

• OLM series of gauges and accelerometers fitted on 2 aircraft

→ With OEM develop maneuver and dynamic FEM

• Apply representative fatigue load spectra and look for new critical locations. Inspect fleet (FALKOR)

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Enhanced Teardown of a PC-9/A Wing (2018)

Program Aims were twofold:

1. to provide experimental evidence mitigating a longstanding RAAF fleet safety risk arising from 142 wing main spar misdrills in 42 RAAF PC-9/A aircraft and;

2. to demonstrate a rapid, novel, full scale structural experimentation technique for in-service management.

Testing completed at 13.6 Hz first bending resonant frequency with failure inside the test section after only 50 cumulative test hours.

In-service DTA re-ran for test loading, good correlation with test.

Surrogate DTA initial flaw size (1.27mm) ok for a missdrill

DTA providing in-service coverage

2 mm

2 mm

Up

Aft

PC-9/A wing main spar enhanced teardown project

Lower forward cap, hole 3 outboard of Rib 6

No in-service x-ray misdrill indications

Corrosion and

fretting product

Test induced fatigue

crack growth

Test induced fatigue

crack growth

Likely fretting between

fastener and misdrilled

bore leading to fatigue

crack growth. Needs a

clean to confirm.

Example origins.

Misdrill outline

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AD Unique Lifing toolsADF require validated “best tools” for aircraft lifing & structural

refurbishment program etc

• Tools should address “spectrum effects” (i.e. variable amp. “VA”)

•Test Interpretation Difficulties: Crack Growth Different between:

• Quantitative Fractography vs Conventional LEFM models

• CA and VA

• Conventional tools just not good enough! Thus developed more robust

tools and data.

So:

• Long-term activities at DST designed to help better understand and

predict fatigue crack growth (small cracks, low to med K)

DISCLAIMER: Not yet 100% DST wide method (i.e. broad church…).

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AD Unique Lifing tools1. Emulators e.g.

1. CGAP

2. FracRisk

2. Disruptors

a. Lead Crack Framework

b. The cubic rule

c. The Dblock approach

d. Da Hartman-Schijve Equation variant (sub-set of

NASGRO)

e. Easigro

Note: Emphasis on Failure not Design

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The metal aircraft fatigue problem space

AA7050

specimen;

fatigued then

loaded to

reveal cracks

(dye

penetrate)

1. The growth of cracks is the only measurable (and thus useful in assessing impact on structural integrity) fatigue metric;

2. For production aircraft materials, cracks that will play a role in the fatigue life of a component nucleate from sub-mm surface or near-surface discontinuities at high stress regions (i.e. hotspots);

3. The majority of these cracks commence growing from near-day one of operations (but time dependent damage e.g. corrosion, fretting etc may also play a role);

4. Upwards of two-thirds of the life-of-type is spent in growing a detectable crack (» 1mm long);

5. Thus the physically short-crack at the low ΔK regime is the area of most interest to fleet management; and

6. However, traditionally most data and analysis have been produced using long (> 1mm long) cracks (limitations acknowledged in ASTM E647).

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LEFM (Standard AFGROW is good?)AA7050, Hornet FT55Spectrum, low Kt coupon, 4 stress levels

0

1

2

3

4

5

6

7

8

9

10

0 5000 10000 15000 20000 25000 30000 35000

Flight Hours

Cra

ck

Dep

th (

mm

)

428.9 MPa Ref Stress

396.5 MPa Ref Stress

358.5 MPa Ref Stress

324.1 MPa Ref Stress

AFGROW Predictions

When plotted like this… well? However:

Each

point is 1

block of

crack

growth

from QF

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LEFM (Standard AFGROW)AA7050, FT55Spectrum

• Need large Equivalent Pre-crack Size (EPS) (>0.05mm)

• Shape not correct (NDI implications)

• Not always conservative, not good enough!

NDIThreshold

Each

point is 1

block of

crack

growth

from QF

Sub-surface

initiated

428.9 MPa

396.5 MPa

358.5 MPa

324.1 MPa

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Lead Crack Growth

• Assumes no load shedding or residual

stresses changes

• a constant per block

• (geometry factor) constant per block or

“similar geometry”

• Crack growth commences shortly after intro

into service

• Tearing ignored as only small fraction of total

life

• Production defects nucleate cracks

Shown to apply for a large class of problems

log a

life

log ao

Assumptions for lead (first failure) surface cracks:

Known Points in Crack

Growth Life

a0 = EPS EIFS

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Lead Cracks: Lead to failure from a discontinuityA few In-Service and FSFT results

0.001

0.01

0.1

1

10

100

0 5000 10000 15000 20000 25000 30000 35000

Simulated Flight Hours

Cra

ck d

ep

th (

mm

)

P3C Wing

DSTO Mirage Wing

A7 Wing, 200 hr Block

T37B Wing Steel Strap

F-16 12L/Spar 6 Zone III

F-16 RP-10 Zone III

F4 C/D Wing Skin

FA-18 FT46 Y598 Stub

F/A-18 FT55 Stbd Wing

F/A-18 FT55 Y453 Web Taper

F/A-18 ST16 Y453 Web Taper

Swiss F&W Mirage Wing BH#2

F111 A4 Splice AL2024

F111 A4 Splice D6ac

F111 A4 FFH58

F111 FAS281 FTG

F111 FFH13 In-service crack

F111 SRO2 A8-109 in-service

FT46 Y598 Stub Frame

Aermacchi In-service

CPLT

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The Cubic Rule(Frost and Dugdale model (circa 1958) – AA7050)

Predict lives for same

spectrum at different

stresses.

= slope

a = a0e Life

= 3

0

0.0002

0.0004

0.0006

0.0008

0.001

0.0012

0 50 100 150 200 250 300 350 400 450 500

Reference Stress Level (MPa)

Gro

wth

Slo

pe [l

og

(mm

)/L

ife (

ho

urs

)]FT55 Slopes

FT55 Average Slopes

APOL Slopes

APOL Average Slopes

FT55 cubic data fit

APOL cubic data fity = 9.505E-12x

3

y = 8.080E-12x3

refadt

da

FT55

APOL

428.9 MPa

396.5 MPa

358.5 MPa

324.1 MPa

Ref Stress

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Cubic Prediction – USAF F-4 coupon test data

0.1

1

10

100

0 1000 2000 3000 4000 5000

F4 248MPaF4 207MPaF4 248MPa predicted from 207MPa growth

Flights

𝑎2 = 𝑎02 𝑒𝜎2𝜎1

3

1𝑁

For Kt = 1.0+

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da/dB = CVA (Kv.a.)m

Aim: Find CVA and n(VA) to give best fit through VA data.

Method:

Treat VA data like CA data

Best linear fit through da/dB vs Kv.a plot on a log-log scale.

• da/dBLOCK measured from Q. fractography

• Kv.a = v.aa where v.a is some characteristic stress• Equation integrated for block-by-block prediction once “Constants” determined.

Effective (Characteristic) Block

Approach (EBA) for repeated VA spectra J. Gallagher Mini-Block

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EBA example

da/dB = CVA (K)2

* Slope = 2 (found to be the case for many spectra investigated!)*

Average C1 derived from curve fit. Average C2 can be estimated

and then used to predict lives of spectrum2.

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-22

-21

-20

-19

-18

-17

-16

-15

0 0.5 1 1.5 2 2.5 3 3.5

log Kref (MPam)

log

(d

a/d

n)

- sm

oo

thed

APOL coupon data

APOL trendline

FT55 coupon data

FT55 trendline

5 peak

stresses/spec

used here

Accel/retardation

inherent in data!

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The Hartman-Schijve Variant

da/dN = D[(ΔK – ΔKthr)/√(1-Kmax/A)]α

Where: Kmax is the maximum stress intensity

D is the da/dN v K y-axis intercept at approximately 1 MPa √m

A is a fracture toughness like parameter

approx = 2

Crack growth data from Virkler et al. and computed variability for AA2024-T3. Half-crack length plotted.

(Note computed ΔKthr = 0 also shown)

Long

Crack

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Easigro

Easigro is a tool which:

• Contains many CG models (including Hartman-Schijve)

• Design to fit optimised curves to da/dN data

• Allows the prediction and visualisation of surface

projection marks

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Types of Discontinuities

● Production components have many sources of discontinuities that can cause fatigue cracking

● Machining damage:

● badly drilled holes (e.g. Macchi etc)

● scratches, grooves, burrs, small tears, nicks

● Surface treatments (pickling, anodizing):

● etch pits, sometimes intergranular attack

● Constituent particles (aluminium alloys and steels)

● particles can be already cracked from production

● Porosity in thick aluminium alloy plate and castings

N.B: discontinuity depths mostly small, ≈ 0.01mm

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Types of Discontinuities (II): Examples

Crack in constituent

particle prior to fatigue

loading

machining damage

lap from shot peening constituent particles

Surface coating

Surface

Fatigue Pore

porosity

Lap

Fatigue crack

Peened surface

Machining tear Surface

Fatigue

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Where to start? With EPS!

0

5

10

15

20

25

30

350

.00

11

0.0

01

9

0.0

03

1

0.0

05

2

0.0

08

8

0.0

14

8

0.0

25

0

0.0

42

0

0.0

70

7

0.1

18

8EPS (mm)

Fre

qu

en

cy

Log-normal distribution of the Equivalent Crack Pre-Size (EPS) of the etched coupon specimens - 120 points

0.001

0.01

0.1

1

0.001 0.01 0.1 1

EPS (BPA, mm)

EP

S V

isu

al

(mm

)

Etched

Peened

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Example C130J WFSFT – CW-1

Beam Cap Flange Crack

Beam Cap Radius Crack (Half Crack)

Panel Crack

RST Loads

Stringer 24

STBD

UNCLASSIFIED – Approved for Public Release

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C130J WFT Fracto & Lead Crack

UNCLASSIFIED – Approved for Public Release

7.000

LC CW-1ALC CW-1

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C130J WFT Fracto & Lead CrackAnalysis Sanity Check

UNCLASSIFIED – Approved for Public Release

7.000

LC CW-1ALC CW-1

28 Achieve game-changing improvements in the way air platforms are sustained

Future ASI for iSustainment

PRESENT STATE• Rules based• Labour-intensive inspections

and analyses• Infrastructure-intensive:

expensive, large, long-running tests for certification and V&V.

• Reactive: based on post-mortem of accidents, incidents & shortfalls

FUTURE STATE• Risk based: probabilistic approaches • Computation intensive, real time material &

structural assessment.• Rapid & flexible design, simulation and

verification approaches• Composite lifing• Certified AM• Non-contact, wide area inspection

AUTOMATION AND ADVANCED SENSING

BIG DATA, AI, HIGH SPEED

COMPUTING, SIMULATION

PROGNOSTICS, 3D PRINTING

Virtual Air System

High Speed Testing

Future sensing & inspections

Future Vertical Lift

Goal: Create a game-changing impact on force delivery & sustainment

Unclassified

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Conclusions

• Adopt a Defence Scientist: how can you make a

difference?

• Look for low hanging fruit

• DASA and AD have a long history of synergistic

achievements

• The road to flight is through the regulator

• Things can always be done better

iSustainment

• Strive for robust tools and methods

• Collaborate everything

• Exciting ASI opportunities just

around the corner

Adopt

me!

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Refs: Kai Maxfield, Matthew McCoy, Douglas Williams, Robert Ogden, Vui Tung Mau and Anthony

Zammit, Failure analyses of a military Transport Aircraft fatigue test, proc. ASIP 2018. Phoenix, AR, USA.

Rudd, JL and Gray, TD, Quantification of Fastener- Hole Quality, Aircraft 1978; 15, 3: 143-147

Virkler DA, Hillberry BM and Goel PK., The statistical nature of fatigue crack propagation. Technical Report AFFDL-TR-78-43, 1978, USA: Air Force Wright Aeronautical Laboratory, Ohio

Molent L, Barter SA and Wanhill RJH. The Lead Crack Fatigue Lifing Framework, Int Fatigue; 33 (2011) 323–331

Jones R, Molent L., Walker K., Fatigue crack growth in a diverse range of materials, Int J Fatigue 2012; 40: 43-50

Gallagher JP and Molent L. The equivalence of EPS and EIFS based on the same crack growth life data, Fatigue 2015; 80:162-170

Gallagher JP. Estimating fatigue-crack lives for aircraft: techniques, Experimental Mech. 1976: 425-433

White P, A guide to the program easigro for generating optimised fatigue crack growth models, DST-Group-TR-3566, DST, Feb 2019

Main, B., Muller, K., Konak, M., Sudhakar, S., Jones, M. and Barter, S. (2019) Evaluation of a PC-9/A Wing Main Spar with Misdrills using Enhanced Teardown, In: Proceedings of the 29th International Conference on Aeronautical Fatigue and Structural Integrity (ICAF) Krakow, Poland

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Questions?

No Highway in the Sky