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)
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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)
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•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
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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
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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
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
AM materials now characterised include:
Ti-6Al-4V
316L Stainless Steel
AerMet 100
Inconel 625
Inconel 718
Al-10Si-0.4Mg
7
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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
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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
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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.
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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.
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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.
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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.
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CUTING THE GORDIAN KNOT
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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
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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.
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EXAMPLE: Ti64 Replacement Part – US Navy P3C Flight
Load Spectrum (FCA 351), Peak Stress 171 MPa
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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.
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LIFING ANALYSIS
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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].
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• 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
Effect of variations in the fracture toughness term
(A) on the computed operational life
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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
.
Effect of variations in the fracture toughness term
(A) on the computed operational life
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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
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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
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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
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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
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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
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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
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• 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
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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
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*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
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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
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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
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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
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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
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QUESTIONS ?????
37 | Monash/RUAG | 12/04/2019
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
• 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
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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
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• 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
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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
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• 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
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All specimens were cut from the panels supplied by MPSPO- Adam
Bowler and Trent Simcock.
43 | Monash/RUAG | 12/04/2019
Test Specimen Geometries
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Specimens with Simulated Skin Corrosion – Cut from P3C wing
panels
44 | Monash/RUAG | 12/04/2019
Test Specimen Geometries
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Specimens with both simulated corrosion and an SPD repair.
45 |
44 | Monash/RUAG | 12/04/2019
Test Specimen Geometries
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
Test Specimen Geometries
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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 -
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
47 | Monash/RUAG | 12/04/2019
Test Specimens –
Courtesy of Adam Bowler and Trent Simcock
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
Effect of Skin Corrosion on Skin Stress
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CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
Stress field on the upper surface of the repaired P3C panel
Skin Stress Distribution is Restored by SPD
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CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
Effect of Skin Corrosion on Buckling Load
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Pristine specimen
Specimen with skin
corrosion
Cold Spray (SPD)
Essentially restored
Buckling modes and
Buckling Modes
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
Effect of SPD on Skin Stresses
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-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
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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!
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CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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, 1.375 Left hand aft upper spar WS346
1 crack, 2.25 Left hand aft upper spar WS584
1 crack, 2.5 Left hand aft lower spar WS349
1 crack, 0.375 Left hand aft lower spar WS380
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CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
• 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.
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CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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.
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54| Monash/RUAG | 12/04/2019
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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!
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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
CRC for Aircraft Airworthiness and SustainmentCoE-Structural Mechanics
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.
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