CRC for Aircraft Airworthiness and Sustainment Analysis of Crack Growth … · 2019. 11. 24. · Inconel 625 Inconel 718 Al-10Si-0.4Mg 7. CRC for Aircraft Airworthiness and Sustainment

CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics

1

Analysis of Crack Growth in AM Replacement Parts

& Laser Additive Repairs

ADF Aircraft Structural Integrity Symposium,

Defence Plaza, Melbourne, 19th – 20th March 2019.

Neil MatthewsSenior Manager, Advanced Technology & Engineering Solutions

RUAG Australia, 836 Mountain Highway, Bayswater, VIC 3153, Australia

&

Professor Rhys Jones, ACCompanion of the Order of Australia

Outline of recent work performed in conjunction with Nam Phan

(NAVAIR) and John Michopoulos (NRL)

US Navy ONR NICOP Grant (N00014-18-S-B001)


2

NAVAIR AM Overview April 3, 2017

Presented To:

Sea Air Space 2017 Presented By:

Ms. Elizabeth McMichael and Dr. William Frazier NAVAIR AM/DT IPT

Background

Slides courtesy of Nam Phan (NAVAIR)


•DISTRIBUTION A. Approved for public release: distribution unlimited. SPR#2017-58 11•DISTRIBUTION A. Approved for public release: distribution unlimited. SPR#2017-58 7

Linking AM to NAVAIR Imperatives

Slide courtesy of Nam Phan (NAVAIR)

•1. Readiness • Parts on Demand

•Distributed Supply Chain

•Local Repair

•2. Increased Speed to the Fleet

•Small, Empowered Teams

•Better Requirements Informed by Experimentation

•Prototyping and Experimentation at all Levels

•Understanding and Acceptance of Appropriate Risk


USAF Vision - Future Role of Structural AM, From: Robert Bair, Chief, Structures Branch, AFAFLCMC/EZFS, USAF.

AM for Structural Parts will realize its greatest potential when it moves beyond non Safety-of-Flight parts

- Cracking / damage of Safety-of-Flight parts is direct driver for part replacement - Will allow for use of AM in a production environment for structurally

significant parts - Speed Repair / Mod lines for on the demand Safety-of-Flight parts

Allow for repair parts outside the supply chain

We must address the “tough” structural certification requirements for AM to realize its full potential.

Structural Certification of Safety-of Flight AM parts means going after the “high hanging fruit”

- Full material allowables development (Strength and DaDT) - Building block approach and scale up testing for parts - Development of analytical methods to analyze AM parts, predict behavior, and minimize future testing

4


The recent papers by the NAVAIR team [1-5] have shown an ability to characterise crack growth in a range of AM materials, as well as in Direct Metal Deposition Repairs, Cold Spray, also known as Supersonic Particle Deposition (SPD), using the Hartman-Schijve (HS) variant of the NASGRO crack growth equation.

1. Jones R., Michopoulos JG., Iliopoulos AP., Singh Raman RK., Phan N. and Nguyen T., Representing Crack Growth In Additively Manufactured Ti-6Al-4V, (2018) International Journal of Fatigue, 2018, 111, pp. 610-622.

2. Iliopoulos AP., R. Jones R., Michopoulos JG., Phan N., Singh Raman RK., Crack growth in a range of additively manufactured aerospace structural materials, Special Issue, Civil and Military Airworthiness: Recent Developments and Challenges, Aerospace, doi:10.3390/aerospace5040118

(On line 9th Dec, ~400 reads to date {last Monday}.)

3. Jones R., Singh Raman RK., Iliopoulos AP., Michopoulos JG., Phan N. and Peng D., Additively manufactured Ti-6Al-4V replacement parts for military aircraft, Int. Journal of Fatigue, https://doi.org/10.1016/j.ijfatigue.2019.02.041

4. Alison J. McMillan, Daren Peng, Rhys Jones, Nam Phan, John G. Michopoulos, Additive manufacturing: Implications of surface finish on component life, Proceedings, SAMPE Europe Conference 2018, Southampton, 11-13th September, 2018.

5. Jones R., Matthews N., Baker A., Champagne V, Aircraft Sustainment and Repair, Butterworth-Heinemann Press, 2018, ISBN 9780081005408. (Book).

5

WHAT HAVE WE DONE AND WHERE DO WE STAND


Crack growth in AM materials can be analysed with existing Damage Tolerance tools.

Crack growth in AM materials conforms to the Hartman-Schijve (HS) equation [1, 2], see Slide 8 for details:

This is true regardless of whether the AM process is:

• Electron beam melting (EBM),

• Direct metal laser sintering (DMLS),

• Selective laser melt (SLM),

• Hot isostatic pressing (HIP),

• Laser engineered net shaping (LENS)

• Whether the LENS process is low- or high-power;

• Whether the build direction was horizontal or vertical

6

CRACK GROWTH IN ADDITIVELY MANUFACTUREDMATERIALS


AM materials now characterised include:

Ti-6Al-4V

316L Stainless Steel

AerMet 100

Inconel 625

Inconel 718

Al-10Si-0.4Mg

7


We have established that crack growth in AM materials can be

expressed as per the HS variant of the NASGRO equation, [1-5]:

Here D and p are material constants, A is the cyclic fracture

toughness, ∆K is the range of the stress intensity factor (K) seen

in a cycle, and ∆Kthr is the associated cyclic fatigue threshold.

For Ti-6Al-4V p = 2.13, D = 2.79 10-10 [1, 2] and

∆Kth ~ ∆Kthr + 0.62

8

da/dN = D (K - Kthr)p/(1-Kmax/A)p/2 (1)

THE HARTMAN-SHIJVE (HS) VARIANT OF THE NASGRO CRACK GROWTH EQUATION


Our 2018 IJF paper [1]

contains approximately

34 examples.

The subsequent 2018

Aerospace paper [2]

contains another 15+

9

Measured and computed long crack da/dN versus ∆K curves for crack growth perpendicular

to the build for SLM Ti-6Al-4V and HIPed SLM Ti-6Al-4V from:Leuders S., Thöne M., Riemer A., Niendorf T., Tröster T., Richard HA., Maier HJ., (2013) On the

mechanical behaviour of titanium alloy TiAl6V4 manufactured by selective laser melting: Fatigue

resistance and crack growth performance, International Journal of Fatigue, 48, pp. 300–307.

SLM = Selective Laser Melt

HT= Heat treated

EXAMPLES


Zhai Y., Galarraga H., Lados DA., (2016) Microstructure, static properties, and fatigue

crack growth mechanisms in Ti-6Al-4V fabricated by additive manufacturing: LENS and

EBM, Engineering Failure Analysis, Vol. 69, pp. 3-14. 10

Comparison of measured

and computed crack

growth for EBM Ti64

specimens with different

build directions.

Note: The similarity of the

these da/dN v ∆K curves to

the bridge (mild) steel curve.

LENS = Laser Engineered

Net Surface

EBM = Electron Beam Melt

EXAMPLESCONTINUED

1.0E-11

1.0E-10

1.0E-09

1.0E-08

1.0E-07

1.0E-06

1.0E-05

1.0E-04

1 10 100 1000

da

/dN

(m

/cyc

le)

ΔK (MPa√m)

AF HOR Computed AF HOR

HT HOR Computed HT HOR

AF VERT Computed AF VERT

HT VERT Computer HT VERT

Bridge steel master curve


Zhai Y., Galarraga H., Lados DA., (2016) Microstructure, static properties, and fatigue

crack growth mechanisms in Ti-6Al-4V fabricated by additive manufacturing: LENS and

EBM, Engineering Failure Analysis, Vol. 69, pp. 3-14. 11

Comparison of measured

and computed crack growth

for the LENS specimens

fabricated with different

laser powers, different heat

treatments and different

build directions.

Note: Similarity of the AM

Ti-6Al-4V da/dN v ∆K curves

to the bridge (mild) steel

Master curve.

EXAMPLESCONTINUED

1.0E-11

1.0E-10

1.0E-09

1.0E-08

1.0E-07

1.0E-06

1.0E-05

1.0E-04

1 10 100 1000d

a/d

N (

m/c

yc

le)

ΔK (MPa√m)

LP AF HOR Computed LP AF HOR

LP HT HOR Computed LP HT HOR

LP AF VERT Computed LP AF VERT

LP HT VERT Computed LP HT VERT

HP AF HOR Computed HP AF HOR

HP HT HOR Computed HP HT HOR

HP AF VERT Computed HP AF VERT

HP HT VERT Computed HP AGED VERT

Bridge steel master curve


12

We have established [1,2] that:

When compared to the variation in the fatigue threshold ∆Kthr seen

for conventionally manufactured materials the variation in the

fatigue threshold ∆Kthr in AM materials is not significantly

(statistically) different.

In contrast, we have established [1, 2] that the AM process can result in a significant variability in the apparent cyclic toughness A.

THE CONTROLLING PARAMETERS & VARIABILITY IN CRACK GROWTH IN AM MATERIALS.


The 2018 State of The Art Review paper [6], [7], the 2018 text [8], and

a large number of publications has established that:

The da/dN versus ΔK curve associated with small naturally occurring

cracks that arise and grow in service can be often be determined

from the long crack HS variant of the NASGRO equation,

representation by setting ΔKthr to be a small value.

6. Jones R., Singh Raman RK., McMillan AJ., (2018) Crack growth: Does

microstructure play a role?, Engineering Fracture Mechanics, 187, pp.

190-210.

7. Tamboli D., Barter S., Jones R., (2018) On the growth of cracks from etch

pits and the scatter associated with them under a miniTWIST spectrum,

International Journal of Fatigue, 109, pp. 10-16.

8. Jones R., Matthews N., Baker AA., Champagne V., Aircraft Sustainment

and Repair, Butterworth-Heinemann Press, 2018, ISBN 9780081005408.13

NATURALLY OCCURRING CRACKS THAT ARISE AND GROW IN SERVICE AIRCRAFT.


We [2] subsequently

confirmed that the

growth of small cracks

in AM Ti64 is captured

by setting ΔKthr to a

small value, typically

(0.1 MPa √m).

Note: The similarity of

the small/short AM

Ti-6Al-4V da/dN v ∆K

curve to the small crack

locomotive and bridge

steels curves.

14

NOTE: Experimental data (AM Ti6Al4V, two different Ti alloys and 2

different mild steels, at a range of R values) reveals that yield stress and

microstructure play little role in the growth of small cracks, see [6, 8].

1.0E-10

1.0E-09

1.0E-08

1.0E-07

1.0E-06

1.0E-05

1 10 100

da/d

N (

m/c

ycle

)

(K) (MPa √m)L1 R = 0.14 L2 R = 0.5L3 R = 0.5 L4 R = -1L5 R = 0.14 L6 R = 0.5L7 R = 0.14 L8 R = 0.14L9 R = 0.5 S11 R = 0.5S12 R = 0.14 S13 R = 0.5Predicted small crack LENS Ti-6AL-4V Small crack Lens Ti6AL4VSmall crack Ti-6Al-4V MA Ti-17 (All R ratio's)USAF Ti-642 R = 0.5 USAF Ti-642 R = 0.05Small crack 1960's Bridge Steel

Specimens L1-L9, S11-S13 are a 350 MPa locomotive mild steel

GROWTH OF SMALL CRACKS IN AM Ti64 COMPARED WITH SMALL CRACKS IN A RANGE OF MATERIALS.


Logistics driven – a part may take more than a year to

arrive.

Question to be answered if AM replacement parts to be

considered:

Will an AM replacement part last the required number of

flight hours?

15

AM REPLACEMENT PARTS


The lifing of AM parts can require the analysis of both

surface roughness AND surface breaking

DISCONTINUITIES that emanate from rough surfaces.

Or (near) subsurface cracks that initiate from porosity,

etc.

This raises the conundrum:

For a limited number of replacement parts, will you

know (need to model) the nature of the

surface/subsurface discontinuities?

Will you know (need to model) the precise nature of the

surface roughness?

16

SURFACE ROUGHNESS AND MATERIAL DISCONTINUITIES


To cut through this “Gordian knot”1 we used, as is commonly

done for conventionally manufactured materials, an Equivalent

Initial Flaw Size (EIFS) approach, see [3].

1The Gordian Knot is a legend of Phrygian Gordium that is

associated with Alexander the Great. An oracle had declared

that any man who could unravel its elaborate knots was

destined to become the ruler of all of Asia. Alexander reasoned

that it would make no difference how the knot was loosed, so

he drew his sword and sliced it in half with a single stroke.

This explanation is taken from Wikipedia.

17

CUTING THE GORDIAN KNOT


Choice of an Equivalent Initial Flaw Size (EIFS).

• To be conservative, and consistent with current practices, we

assumed an EIFS of 1.27 mm (0.05 inch).

• USAF DTA requirement.

• This is significantly bigger than surface roughness and also typical

AM defects, lack of fusion, see [3, 9].

• Consequently, a precise knowledge of surface roughness and

material discontinuities is not needed.

9. Romano S., Brandão A., Gumpinger J., Gschweitl M., Beretta S.,

Qualification of AM parts: Extreme value statistics applied to tomographic

measurements, Materials and Design, 131, pp 32-48, 2017.18

CUTING THE GORDIAN KNOT - CONTINUED


The specimen geometry analysed was an 80 mm wide by 2.6 mm

thick AM Ti6Al-4V wing skin specimen containing a centrally

located 6 mm diameter hole. (Represents a fastener hole in a P3C

Orion wing skin.)

Fatigue critical location FCA-351 is in the P3C wing skin near the

lower front spar inboard.

19

EXAMPLE: Ti64 Replacement Part – US Navy P3C Flight

Load Spectrum (FCA 351), Peak Stress 171 MPa


Used the small crack growth equation for small naturally

occurring cracks in AM Ti-6Al-4V, given in [2]:

da/dN = 2.79 x 10-10 [(K - 0.1) /(1-Kmax/A)1/2]2.13

This equation is shown in Slide 14 together with the

measured small crack growth curve.

20

LIFING ANALYSIS


A (MPa √m)

Cyclic Fracture

Toughness

Computed

operational life

(Simulated flight

hours)

Estimated operational

life, i.e. the computed

life/3.

(Simulated flight hours)

128 (Conventional material)

19418 6473

75 (Mean value [1, 2]) 17075 5691

62 15839 5279

36.6 (Boeing Data) 9446 3148 (Not good enough)

A = 62 MPa √m,

EIFS = 0.69 mm

18300 6133

A = 62 MPa √m,

EIFS = 0.448 mm(AIRBUS Data)

20436 6812

21

Effect of variations in the fracture toughness term

(A) on the computed operational life, from [3].


• Design life of P3C is 15,000 flight hrs

• So for an EIFS of 1.27 mm the previous slide

reveals that if the life requirement for the part was

5000 flight hours, then only those AM processes

with a fracture toughness of > 62 MPa √m would be

acceptable.

• Assuming an EIFS of 0.69 mm, which is still bigger

than those determined by AIRBUS for AM Ti64, the

life of the part increases by approximately 900 flight

hrs to approximately 6100 flight hrs, from [3].

22


(A) on the computed operational life


23

0.1

1

10

100

0 5,000 10,000 15,000 20,000 25,000

Cra

ck d

ep

th (

mm

)

Flt hrs

A = 128

A = 75

A = 62

A = 36.6

EIFS = 0.69 mm, A = 62

EIFS = 0.448 mm, A = 62

.


(A) on the computed operational life


Crack Growth in Laser Additive Repairs

Example: Consider the crack growth data presented in [10] for the

growth of surface breaking cracks in a 20 mm diameter round bar of

AerMet100 subjected to variable amplitude loading. These specimens

had an initial, approximately 0.25 mm deep, notch.

10. Walker KF., Lourenço JM., Sun S., Brandt M., Wang CH., Quantitative

fractography and modelling of fatigue crack propagation in high strength

AerMet100 steel repaired with a laser cladding process, International Journal of

Fatigue, 94 (2017), pp. 288–301. 23


Load Spectrum - From [9]. σmax σmin

No of

Cycles

1000 100 1000

1000 700 500

1000 100 10

1000 700 500

1000 100 10

1000 700 500

1000 100 500

1000 700 500

1000 100 10

1000 700 500

1000 100 10

1000 700 500

1000 100 10

1000 700 500

1000 100 500

1000 700 500

1000 100 10

1000 700 500

1000 100 10

1000 700 500

1000 100 10

1000 700 500

1000 100 10

1000 700 500

1000 100 500

1000 100 10024


Both baseline damaged, and LAD repaired

specimens tests were analysed

LAD repaired specimen, from [10]

Baseline with a 0.25 mm deep

starter crack in an AerMet100 steel

specimen, from [10].

25


Predicting Crack Growth

The Hartman-Schijve (HS) variant of the NASGRO crack growth equation

for AerMet100 was given in [2] as:

da/dN = 5.06 10-10 [(K - Kthr) /(1-Kmax/A)1/2]1.81 (2)

In our analysis we used the value of A given in [10], viz: A = 140 MPa √m.

Hence, all that is needed in order to use the Equation (2) to compute the

crack growth history for these tests is the value of Kthr.

26


Predicting Crack Growth

Since the initial crack is small we used a small value for the threshold

term Kthr , see [2, 6-8, 11-13].

The present analysis used the lower bound value Kthr = 0.1 MPa √m

that is recommended in [6-8, 11-13] when computing the growth of

small naturally occurring cracks.

11. Jones R., Fatigue crack growth and damage tolerance, Fat Fract Eng Mat and Struct, 2014;

37, pp. 463–483.

12. Jones R., Peng D., McMillan A., Crack growth from naturally occurring material

discontinuities, Chapter 5, pp. 129-190, Aircraft Sustainment and Repair, Edited by R.

Jones, N. Matthews, AA. Baker and V. Champagne Jr., Butterworth-Heinemann Press,

2018, ISBN 9780081005408.

13. Main B., Evans R., Walker K., Yu X., Molent L., Lessons from a fatigue prediction challenge

for an aircraft wing shear tie post, International Journal of Fatigue, 123, pp. 53-65, June

2019.

27


Comparison Of Predictions For The Baseline Tests

Excellent agreement between computed and measured crack

growth histories given in Walker et al. (DST and RMIT) [9].

0.10

1.00

10.00

0 5000 10000 15000 20000 25000

Cra

ck d

epth

Cycles

Specimen 14

Specimen 15

Computed Walker et al

Computed

30


• We have established an ability to account for the variability in

crack growth.

• This is a fundamental requirement of MIL-STD-1530.

• As shown in [1, 2, 6-8, 11, 12, 14, 15, etc] this can be done by

allowing for variability in the threshold term Kthr.

[14] Jones R., Kinloch AJ., Michopoulos JG., Brunner AJ. and Phan N., (2017)

Delamination growth in polymer-matrix fibre composites and the use of

fracture mechanics data for material characterisation and life prediction,

Composite Structures, 180, pp. 316-333.

[15] Yao L., Alderliesten R., Jones R., Kinloch AJ., (2018) Delamination Fatigue

Growth in Polymer-Matrix Fibre Composites: A Methodology for Determining

the Design and Lifing Allowables, Composite Structures, 96, pp. 8-20.29

Computing The Variability In Crack Growth


Comparison of predictions, and test data for the

LAD repaired specimens – variability is captured

0.10

1.00

10.00

0 25000 50000 75000 100000 125000 150000

Cra

ck d

epth

(m

m)

Cycles

Specimen 1Specimen 2Specimen 3Computed Specimen 1Computed Specimen 2Computed Specimen 3Computed Walker et alComputed Mean-3σ

HAZ

30


*Used the Mean - 3σ value to determine the fastest possible

curve, as required by MIL-STD-1530

Specimen Kthr (MPa √m) A (MPa √m)

1 7.0 140

2 9.0 140

3 5.8 140

Mean – 3σ* 0.19 140

33

Threshold Values Used To Capture Variability


CONCLUSIONS

• MIL-STD-1530 requires an ability to assess the variability in life

and thereby the fastest possible growth, i.e. minimum life.

• In this context we have an ability to capture the variability in

the fatigue lives associated with post heat treated laser clad

specimens.

• At first glance, the experimental data suggests that the fatigue

behaviour of post heat treated clad Specimens 1-3 would

appear to be superior to that of the baseline AerMet100 steel.

• However, the variability in the fatigue lives is such that this

“apparent” superior performance should not be taken into

account when assessing the fatigue performance of post heat

treated laser clad specimens.

33


One of the initial challenges to be faced by additively

manufactured parts is logistics related.

By this we mean that:

The long time scales that can be associated with the

procurement, and the availability of a conventionally

manufactured part may mean that to maintain aircraft

availability the use of an AM part that has an operational

life that is less than the original design life of the

airframe may be an attractive option.

34

DISCUSSION


To achieve this MIL-STD-1530D notes that a Damage

Tolerant Analysis capability is essential.

To this end we have developed the capability to accurately

compute the growth of cracks in:

AM materials

LAD repairs

Cold spray (SPD) repairs

& CAPTURE THE VARIABILITY IN MATERIAL BEHAVIOR:

A KEY requirement of MIL-STD-1530D.

35

SUMMARY


The Damage Tolerant analysis tools developed for

Conventionally Manufactured Materials also hold for AM &

for Repairs using Laser Deposition and Cold Spray (SPD).

36

1.0E-10

1.0E-09

1.0E-08

1.0E-07

1.0E-06

1.0E-05

1 10 100

da/d

N (

m/c

ycle

)

(K) (MPa √m)L1 R = 0.14 L2 R = 0.5L3 R = 0.5 L4 R = -1L5 R = 0.14 L6 R = 0.5L7 R = 0.14 L8 R = 0.14L9 R = 0.5 S11 R = 0.5S12 R = 0.14 S13 R = 0.5Predicted small crack LENS Ti-6AL-4V Small crack Lens Ti6AL4VSmall crack Ti-6Al-4V MA Ti-17 (All R ratio's)USAF Ti-642 R = 0.5 USAF Ti-642 R = 0.05Small crack 1960's Bridge Steel

Specimens L1-L9, S11-S13 are a 350 MPa locomotive mild steel

Bottom Line


QUESTIONS ?????

37 | Monash/RUAG | 12/04/2019


• We have also shown how SPD (Cold Spray)

can be used to alleviate the effect of:

• SCC in risers.

• Intergranular cracking at fastener holes.

• Skin corrosion.

38

RELATED SLIDES – FOR FURTHER REFERENCE


Supersonic Particle Deposition (SPD) (Cold Spray)

A technology in which metal powder particles are injected into

a supersonic gas flow and impact a solid surface with

sufficient energy to cause plastic deformation and bonding

with the underlying material.

Bonding is a result of high strain rate deformation and

adiabatic shear instabilities and the bond interface.

No heat affected zone, no interface oxides, generation of

surface compressive stresses, no thickness limitations.

Various powder depositions (Aluminium, Nickel, Titanium,

Inconel, Steel) on various aerospace metal substrates.

Mg

Al

39 | Monash/RUAG | 12/04/201938


• Surface/Skin corrosion often presents as exfoliation or surface

pitting or a combination of both. In many instances, corrosion

initiates due to the presence of dissimilar metals (e.g.

fasteners/skins) and adverse environments (e.g. salt water).

• The problem of skin corrosion

is seen in many areas of the

P3C and often results in

multiple co-located repairs.

Parasitic stiffening;

Changes in load path;

40 | RUAG Australia | 12/04/2019Multiple corrosion repairs on a P3C aircraft wing,

(Courtesy of MPSPO.)

Skin Corrosion


Assessment of Wing Integrity with Upper Surface

Corrosion Subject to Compressive Loading

MPSPO and AGAP advised that typical

surface corrosion often has a near

circular planform.

When assessing the impact of corrosion

a full width corrosion grindout is

normally assumed.

This section presents an evaluation of

the effect of skin corrosion which is

removed leaving a full width grindout.

In undertaking the analysis cognizance

has been taken of the US Joint Services

Structural Guidelines JSSG2206 which

specifies that there must be no yielding

at 115% Design Limit Load (DLL). 41 | Monash/RUAG | 12/04/2019


• The geometry of the panels and risers analysed was taken so

as to represent a typical AP3C wing section and more

particularly panels supplied by MPSPO.

View of the interior of a P3C wing showing the location of the H-clips

and a section of the wing . (Pictures courtesy of MPSPO).

Assessment of Wing Integrity with Upper Surface

Corrosion Subject to Compressive Loading

42 | Monash/RUAG | 12/04/2019


All specimens were cut from the panels supplied by MPSPO- Adam

Bowler and Trent Simcock.

43 | Monash/RUAG | 12/04/2019

Test Specimen Geometries


Specimens with Simulated Skin Corrosion – Cut from P3C wing

panels

44 | Monash/RUAG | 12/04/2019



Specimens with both simulated corrosion and an SPD repair.

45 |

44 | Monash/RUAG | 12/04/2019




46 | Monash/RUAG | 12/04/2019

Dimensions

Specimen Number

1# 2# 3# 4# 5# 6#

(mm) (mm) (mm) (mm) (mm) (mm)

B 4.27 4.37 3.97 4.07 3.95 3.98

H 32.80 32.90 30.70 30.90 30.60 31.20

D 122.40 122.00 113.00 114.50 113.90 113.70

D1 4.00 3.00 1.50 2.00 2.00 1.50

D2 4.50 4.50 2.30 2.00 0.50 1.00

TB 2.26 2.72 2.25 2.32 2.22 2.46

Tc - 0.95 0.93 1.38 1.05 -

Ts - - 2.73 - 2.42 -


47 | Monash/RUAG | 12/04/2019

Test Specimens –

Courtesy of Adam Bowler and Trent Simcock


Effect of Skin Corrosion on Skin Stress

48 | Monash/RUAG | 12/04/2019


Stress field on the upper surface of the repaired P3C panel

Skin Stress Distribution is Restored by SPD

49 | Monash/RUAG | 12/04/2019


Effect of Skin Corrosion on Buckling Load

50 | Monash/RUAG | 12/04/2019

Pristine specimen

Specimen with skin

corrosion

Cold Spray (SPD)

Essentially restored

Buckling modes and

Buckling Modes


Effect of SPD on Skin Stresses

51 | Monash/RUAG | 12/04/2019

-420.0

-370.0

-320.0

-270.0

-220.0

-170.0

-120.0

-70.0

-20.0

-210.0 -180.0 -150.0 -120.0 -90.0 -60.0 -30.0 0.0

Str

ess

(M

Pa)

Axial Force (kN)

Strain Gauge 1 Strain Gauge 2

Strain Gauge 3 Strain Gauge 4

357 MPa

-183 kN

Panel stress at

Limit load ~ 170

MPa.

SPD repaired

panel can

withstand proof

load, with no non

linear behaviour at

115% DLL –

A JSSG2006

requirement


SPD Repair To A P3C Wing Plank With

Skin Corrosion Between Stiffeners

The experimental and analysis were in excellent agreement.

Revealed that SPD repairs, for the case when there is up to a

50% loss of material between the risers, can restore the load

carrying capacity of the wing!

52 | Monash/RUAG | 12/04/2019


Stress corrosion cracking (SCC) in rib stiffened wing

planks

• Stress corrosion cracking (SCC) in rib stiffened

wing planks is common to both military

transport and maritime reconnaissance aircraft.

• SCC is also seen in P3C Orion aircraft as is

evident from SCC in RAAF AP3C aircraft A09-

659. Data courtesy James Ayling (AGAP).

Schematic diagram of

location of stress

corrosion cracking in

C-130 wing

Number of SCC cracks and size

(inches)

Location Wing station

2 cracks, 0.250 & 1.5 Left hand aft upper spar WS143

3 cracks, 0.200, 0.250 & 0.875 Left hand aft upper spar WS209

1 crack, 0.875 Left hand aft upper spar WS275



1 crack, 2.5 Left hand aft lower spar WS349

1 crack, 0.375 Left hand aft lower spar WS380

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• Analysis of a 7 stiffener panel with SCC.

SPD Repairs on Compression Surfaces of a

P3C Wing Plank

Buckling load (No SCC) - 946 kN Buckling load (3 risers with SCC) - 510kN

Application of three 0.2 mm thick, 10 mm

wide, and 110 mm long SPD doublers

essentially restored both the buckling load

and the buckling mode.

54 | Monash/RUAG | 12/04/2019


SPD Repairs on Compression Surfaces of a Wing

Plank - Experimental Validation

Nominally identical specimens were cut from a P3C wing plank provided by

MPSPO.

Specimens with no SCC, with 100 mm SCC, and with both SCC and a SPD

repair were tested.

55 |

54| Monash/RUAG | 12/04/2019


SPD Restores Both Load Carrying Capacity and

Buckling Mode

• SPD restored both the load carrying capacity and the load

deflection curve to that of the structure without SCC.

• Furthermore the SPD repair did not change the load path!

56 | Monash/RUAG | 12/04/2019

0

10

20

30

40

50

60

0 1 2 3 4 5

Axia

l F

orc

e (

KN

)

Axial Displacement (mm)

baseline specimen 1: no SCC

with SCC crack

baseline specimen 2: no SCC

SPD repaired specimen


SUMMARY

SPD repairs are viable options to restore structural integrity. This is

established via the following examples:

External skin “patch” repair to skin corrosion

External “patch” repair to SCC in risers

External “patch” repairs to inhibit intergranular cracking (IGC).

The effect of SPD repairs can be accurately modelled and analysed and in

most cases the analyses have been validated via coupon testing or

simulated wing elements.

SPD repairs to skin corrosion on compression surfaces where there is up

to a 50% loss of material between the risers can restore the load carrying

capacity of the wing.

Stress Corrosion Cracking (SCC) in risers: SCC can result in failure due to

local buckling that can run the length of the section. Analysis and validation

testing has shown that that for SCC in the risers an SPD repair can essentially

restore the load carrying capacity of the wing.

57 | Monash/RUAG | 12/04/2019

CRC for Aircraft Airworthiness and Sustainment Analysis of Crack Growth … · 2019. 11. 24. · Inconel 625 Inconel 718 Al-10Si-0.4Mg 7. CRC for Aircraft Airworthiness and Sustainment

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