The Effect of Peening on the Fatigue Life of 7050 Aluminium Alloy

<strong>The</strong> <strong>Effect</strong> <strong>of</strong> <strong>Peening</strong> on the Fatigue Life <strong>of</strong> 7050
Aluminium Alloy
P.K. Sharp and G. Clark
Airframes and Engines Division
Aeronautical and Maritime Research Laboratory
DSTO-RR-0208
ABSTRACT
Many changes in the design and manufacture <strong>of</strong> high-performance military aircraft ⎯ for
example, the use <strong>of</strong> highly optimised design and the use <strong>of</strong> higher-strength material ⎯ have led
to an increased sensitivity <strong>of</strong> airframe fatigue life to surface features such as corrosion or
mechanical damage. <strong>The</strong> peening applied to the F/A-18 represents a significant departure from
traditional manufacture, and it is therefore important that the RAAF and AMRL have a
thorough understanding <strong>of</strong> the peening process, the surface conditions produced, and their
effect on structural integrity.
This report discusses the fatigue crack growth research at AMRL, and elsewhere, relating to
peening <strong>of</strong> aluminium alloys, and summarises the improvements in peening which have arisen
from this research. . <strong>The</strong> overall aim <strong>of</strong> the peening research and development discussed was to
establish a Life-Improvement-Factor (LIF) for the peening used on the F/A-18, as well as any
future peening required by modifications. It also attempted to provide a means <strong>of</strong> measuring
peening quality, to allow the full exploitation <strong>of</strong> peening to improve fatigue life. It also
highlights areas where further research could be beneficial in relation to peening and the
structural integrity <strong>of</strong> the F/A-18 aircraft. <strong>The</strong> report highlights the practical problems <strong>of</strong>
introducing changes to fatigue critical surfaces, with particular reference to the RAAF and CF
fleets.
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<strong>The</strong> <strong>Effect</strong> <strong>of</strong> <strong>Peening</strong> on the Fatigue Life <strong>of</strong> 7050
Aluminium Alloy
Executive Summary
Many changes in the design and manufacture <strong>of</strong> high-performance military aircraft ⎯
for example, the use <strong>of</strong> highly optimised design and the use <strong>of</strong> higher-strength
materials ⎯ have led to an increased sensitivity <strong>of</strong> airframe fatigue life to surface
features such as corrosion or mechanical damage. <strong>The</strong> peening applied to the F/A-18
represents a significant departure from traditional manufacture, and it is therefore
important that the RAAF and AMRL have a thorough understanding <strong>of</strong> the peening
process, the surface conditions produced, and their effect on structural integrity.
To build the required level <strong>of</strong> understanding, AMRL has undertaken a number <strong>of</strong>
research programs, over nearly ten years, investigating the effects <strong>of</strong> glass bead
peening on the aluminium alloy 7050 used in RAAF F/A-18 aircraft. <strong>The</strong> overall aim <strong>of</strong>
the peening research and development discussed was to establish a Life-Improvement-
Factor (LIF) for the peening used on the F/A-18, as well as any future peening required
by modifications. It also attempted to provide a means <strong>of</strong> measuring peening quality,
to allow the full exploitation <strong>of</strong> peening to improve fatigue life.
From this research came methods <strong>of</strong> surface peening and methods <strong>of</strong> surface removal,
which allowed development <strong>of</strong> localised life extension methods for the airframe. <strong>The</strong>se
life extension methods are now being applied in full-scale fatigue tests (to keep tests
going after initial failures) and are likely to be applied to both RAAF and CF fleet
aircraft.
This report discusses the fatigue crack growth research at AMRL, and elsewhere,
relating to peening <strong>of</strong> aluminium alloys, and summarises the improvements in peening
which have arisen from this research. It also highlights areas where further research
could be beneficial in relation to peening and the structural integrity <strong>of</strong> the F/A-18
aircraft. It highlights the practical problems <strong>of</strong> introducing changes to fatigue critical
surfaces, with particular reference to the RAAF and CF fleets.

Authors
P.K. Sharp
Airframes and Engines Division
____________________
Khan Sharp, Research Scientist. Graduated from Monash
University in 1987 having obtained a Materials Engineering
Degree with Honours. In 1990 he completed a Masters <strong>of</strong>
Engineering Science degree, and commenced work in the Fatigue
and Fracture Detection and Assessment area at Fishermans Bend.
Over the past 10 years he has been involved in the metallurgical
investigation <strong>of</strong> aircraft structures and components, fractographic
analysis <strong>of</strong> fatigue surfaces and research into fatigue crack growth
and fracture <strong>of</strong> aircraft materials. During that time he has
published over 40 reports and papers. He has completed extensive
research into novel methods <strong>of</strong> retarding crack growth and
innovative NDI methods. He presently manages the Structural
Implications <strong>of</strong> Corrosion task as well as conducting research on
fatigue and fracture.
________________________________________________
G. Clark
Airframes and Engines Division
Graham Clark, Principal Research Scientist. Graduated from
University <strong>of</strong> Cambridge in 1972 in Natural Sciences. After
completing research for a PhD on the growth <strong>of</strong> fatigue cracks at
notches, he undertook post-doctoral research at Cambridge on the
detection and growth <strong>of</strong> cracks in submarine pressure vessels. In
1977 he commenced work at DSTO in Maribyrnong, leading
research on cracking in thick-walled pressure vessels; which
developed a comprehensive fracture control plan for Australian
manufactured ordnance and a capability for predicting ordnance
fatigue lives. In 1984 he moved to Fishermans Bend, where he
established a research program on the damage tolerance <strong>of</strong> thick
carbon-fibre composite materials, involving modelling and
experimental investigation <strong>of</strong> impact damage in aircraft materials.
In his present position, he leads tasks which support defect
assessment in ADF aircraft, NDI evaluation and fatigue crack
growth research. He is also chairperson <strong>of</strong> the AMRL Accident
Investigation Committee.
____________________
________________________________________________

Contents
1. INTRODUCTION ............................................................................................................... 1
2. REVIEW ........................................................................................................................2
2.1 General........................................................................................................................ 2
1.2 F/A-18 <strong>Peening</strong> Processes ........................................................................................ 6
3. AMRL RESULTS ............................................................................................................... 14
3.1 Overview <strong>of</strong> <strong>Peening</strong> Research Programs .......................................................... 14
3.2 Early research - definition <strong>of</strong> potential problem .............................................. 14
3.2.1 Assessment <strong>of</strong> RAAF aircraft................................................................. 21
3.2.2 Surface roughness/residual stress measurement............................... 22
1.1.3 Summary <strong>of</strong> AMRL <strong>Peening</strong> Specification .......................................... 26
1.3 Rework <strong>Peening</strong> Research..................................................................................... 26
1.3.1 Different rework process comparisons................................................. 27
1.1.2 F/A-18 fleet rework simulation............................................................. 33
1.1.3 Summary <strong>of</strong> AMRL Rework <strong>Peening</strong> Specification............................ 35
1.4 <strong>Peening</strong> Life Improvement Factor (LIF).............................................................. 35
1.5 Fleet Applications Research.................................................................................. 37
1.5.1 Y470 Crotch Repair Program ................................................................. 37
1.5.2 Fleet Y470.5 X19 Pocket Repair program ............................................. 40
1.1.3 IFOSTP Fatigue Tests .............................................................................. 45
4. DISCUSSION..................................................................................................................... 46
5. CONCLUSION................................................................................................................... 48
6. ACKNOWLEDGMENTS ................................................................................................. 52
APPENDIX A: ...................................................................................................................... 53
6.1 Spectrum Details..................................................................................................... 53
6.2 Almen Definition.................................................................................................... 53
6.3 Saturation Curve ..................................................................................................... 54
7.4 Surface Roughness.................................................................................................. 56

DSTO-RR-0208
1. Introduction
During the early 1980’s the Royal Australian Air Force (RAAF), purchased 74 F/A-18 aircraft
from the USN and the Canadian Forces (CF) purchased 124 F/A-18 aircraft from McDonnell
Douglas Aerospace (MDA). <strong>The</strong> RAAF and CF agreed to implement a combined structural
test program, the International-Follow-On-Structural Test Program (IFOSTP) to refine
estimates <strong>of</strong> the F/A-18 service fatigue lives for the RAAF and CF. <strong>The</strong> F/A-18 has a
number <strong>of</strong> structural integrity features that are new to the Australian and Canadian
experience. <strong>The</strong>se features include widespread use <strong>of</strong> cold working <strong>of</strong> fastener holes and ion
vapour deposition (IVD) treatment <strong>of</strong> metallic components. It also features the widespread
use <strong>of</strong> shot peening <strong>of</strong> large aluminium components; this process involves blasting the
surface with metal, glass or ceramic beads, thus deforming the surface and leaving it
containing residual stresses. <strong>The</strong>se stresses are usually beneficial in that the compressive
stresses near the surface retard the growth <strong>of</strong> fatigue cracks which initiate at the surface. <strong>The</strong>
F/A-18 fatigue life is managed on a Safe-Life basis; its service life is based conservatively on
the non-peened surface condition. However, localised areas <strong>of</strong> the main structure are
peened to enhance fatigue life; examples include situations where modifications are made
under an ECP- (Engineering Change Program) or after repairs.
Shot peening treatments are widely used in mechanical and aeronautical engineering to
improve the fatigue performance <strong>of</strong> components. <strong>The</strong> process is normally associated with
high-strength steels and titanium alloys, although in the F/A-18 it is now being applied
widely to high-strength aluminium alloys. <strong>The</strong>re have however been reports <strong>of</strong> extensive
variations in the fatigue life results for peened components and, in some cases (Clayton and
Clark 1988), a decrease in fatigue life has been observed. Such variability naturally raises
concern that the peening process developed for use with steels might not always be suitable
for peening high-strength aluminium alloys.
Many developments in the design and manufacture <strong>of</strong> high-performance military aircraft ⎯
for example, the use <strong>of</strong> highly optimised design and the use <strong>of</strong> higher-strength material ⎯
have led to an increased sensitivity <strong>of</strong> airframe fatigue life to surface features such as
corrosion or mechanical damage. <strong>The</strong> peening applied to the F/A-18 represents a major
change to the aircraft’s surface condition, and it is therefore important that the RAAF have a
thorough understanding <strong>of</strong> the peening process, the surface conditions produced, and their
effect on structural integrity.
To build the required level <strong>of</strong> understanding, AMRL has undertaken a number <strong>of</strong> research
programs, over nearly ten years, investigating the effects <strong>of</strong> glass bead peening on the
aluminium alloy 7050 used in RAAF F/A-18 aircraft. <strong>The</strong> overall aim <strong>of</strong> the peening research
and development discussed was to establish a Life-Improvement-Factor (LIF) for the
peening undertaken on the F/A-18 or any future peening. It also attempted to provide a
means <strong>of</strong> measuring peening quality, and thereby to allow the full exploitation <strong>of</strong> peening to
improve fatigue life.
From this research came methods <strong>of</strong> surface peening and methods <strong>of</strong> surface removal, which
allowed the development <strong>of</strong> a localised life extension procedure for the airframe. This life
1

DSTO-RR-0208
extension method has now being applied in full-scale fatigue tests (to keep component tests
going after initial failures) and may be applied to both RAAF and CF fleet aircraft.
This report discusses the fatigue crack growth research at AMRL, and elsewhere, relating to
peening <strong>of</strong> aluminium alloys, and summarises the improvements in peening which have
arisen from this research. It also highlights areas where further research could be beneficial
in relation to peening and the structural integrity <strong>of</strong> the F/A-18 aircraft. It highlights the
practical problems <strong>of</strong> introducing changes to fatigue critical surfaces, with particular
reference to the RAAF and CF fleets.
2.1 General
2. Review
This section provides a summary <strong>of</strong> research into the peening <strong>of</strong> aluminium alloys and its
effect on fatigue life. <strong>The</strong>re is an extensive array <strong>of</strong> literature dating back to early this
century examining the effects <strong>of</strong> peening on high-strength steels and titanium alloys and this
will not be discussed here.
<strong>The</strong> two manufacturing parameters which are normally defined in peening any component
are, coverage (proportion <strong>of</strong> surface area peened) and saturation (amount <strong>of</strong> peening energy
applied to each area). <strong>The</strong>se factors act together with the applied stress to determine the
effectiveness <strong>of</strong> peening as a method <strong>of</strong> improving fatigue life. <strong>The</strong> process is calibrated in
terms <strong>of</strong> “Almen Intensity” and Appendix 1 provides a description <strong>of</strong> Almen Intensity,
Coverage and Saturation).
<strong>Peening</strong> attempts to spread material near the impact point against the resistance <strong>of</strong>
neighboring material, thus introducing a complex sub-surface residual stress distribution in
which generally, the surface is in elastic compression. <strong>The</strong>re is a rapid transition to elastic
tension at a deeper level, produced by the “enlarged” surface layer. This tension decays in
deeper regions towards zero, Figure 1. <strong>The</strong> effect <strong>of</strong> these stresses acting across the plane <strong>of</strong>
any growing crack arises principally from the compressive stress, which acts to close the
crack, retarding its growth. <strong>The</strong> retardation caused by these compressive stresses still
operates when the crack tip has grown into the tensile stress field in the core, and is only
gradually reduced by the tensile stress component, as the crack grows deeper into the core.
<strong>The</strong> effect <strong>of</strong> the stresses is always to retard the crack growth.
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Peened surface
+ve
-ve
Residual stress field
after peening
Figure 1: Schematic <strong>of</strong> residual stress distribution below a peened surface. With traditional bead
peening techniques this compressive layer generally extends 200-400µm below the surface.
When peening steel components, the fatigue life is influenced principally by the distribution
<strong>of</strong> residual stress (depth and peak <strong>of</strong> the residual compressive stress). Similarly, in
aluminium alloys the depth and peak <strong>of</strong> the residual compressive stress are also major
factors in determining fatigue life; but because <strong>of</strong> the greater damage to the s<strong>of</strong>ter material,
the surface finish (roughness and defects) is also recognised as a critical influence on fatigue
life. In addition to the research discussed later in this report, there are a number <strong>of</strong> papers
examining this effect, perhaps best summarised in ESDU, 1992. Figure 2 clearly shows the
effect <strong>of</strong> Almen intensity on endurance limit for aluminium alloys. In this example, curve 4,
which has the highest peening intensity, while showing improved endurance at high
alternating stresses has a lower endurance at lower stresses compared to the lower intensity
peening.
A significant effect <strong>of</strong> peening intensity on fatigue life has also been noted; there appears to
be an optimum peening intensity to achieve the longest improvement in fatigue life, Figure
3. This intensity varies with the type <strong>of</strong> aluminium and heat treatment (Luo et al., 1986 and
Luo et al., 1987).
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Figure 2: A graph illustrating the effect <strong>of</strong> surface finish on the endurance <strong>of</strong> aluminium material
(ESDU, 1992). Cross - unpeened (curve 1), Open circle - peened [12-16A] (curve 2),
Solid triangle - peened [16-20A] (curve 3), Open triangle - peened [8-10C] (curve 4).
Note a “C” Almen strip is almost twice as thick as an “A” Almen strip, see Appendix 1.
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Figure 3: <strong>The</strong> effect on fatigue endurance <strong>of</strong> Almen peening intensity for 7075-T73 and 7075-T6
(ESDU, 1992).
Mutoh et. al (1987) performed research which showed that crack initiation lives <strong>of</strong> peened
aluminium alloys (7010 and 7050) were shorter than those <strong>of</strong> unpeened specimens, though
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the total life is longer. <strong>The</strong> difference in fatigue lives between the peened and unpeened
specimens was observed to increase as the stress amplitude decreased.
Meguid (1991) also performed research examining the effect <strong>of</strong> partial peening coverage. He
concluded that, in s<strong>of</strong>t material like aluminium alloys, the plastic zone associated with the
peening dent is several times greater than the size <strong>of</strong> the dent. <strong>The</strong>refore, even a partially
peened surface can show an improved fatigue life over a non-peened surface. In contrast,
however, this is not the case with partially peened high strength steel specimens where the
plastic zone is <strong>of</strong> a similar size to the peening indentation.
In the early 1980’s, during an overhaul <strong>of</strong> an USAF C-141 aircraft, an investigation was made
<strong>of</strong> the peened surfaces. It was observed that at the intensity levels within those prescribed by
MIL-STD-852 the 7075-T6 peened surfaces exhibited peened surface extrusion folds (PSEF)
and numerous laps. Simpson (1985) noted that the depth <strong>of</strong> these laps was a high percentage
<strong>of</strong> the total plastic deformation depth and drew the following conclusions from a parallel
coupon test program.
1) multiples <strong>of</strong> 100% saturation commonly specified in Aerospace peening specifications
may be deleterious to fatigue life in some materials, and
2) fatigue life and peening intensity had a positive correlation.
<strong>The</strong> results outlined above are representative <strong>of</strong> research worldwide which indicates the
importance <strong>of</strong> process quality control in peening aluminium alloys in terms <strong>of</strong> achieving the
optimum improvement in fatigue life. This is not a simple problem and a number <strong>of</strong>
parameters have to be carefully controlled to maximise fatigue life. <strong>The</strong>se include media
quality, Almen intensity, and coverage. As a result <strong>of</strong> concerns about the sensitivity <strong>of</strong>
fatigue life <strong>of</strong> aluminium alloys to surface condition, AMRL initiated research into peening,
to enable the RAAF and its contractors to optimise peening conditions to achieve the best
possible fatigue life improvement.
2.2 F/A-18 <strong>Peening</strong> Processes
<strong>The</strong> majority <strong>of</strong> the research into peening <strong>of</strong> aluminium alloys conducted by AMRL relates to
the lifing <strong>of</strong> the FS488 and FS470.5 wing carry-through bulkheads. <strong>The</strong> research, however, is
sufficiently generic as to be useable elsewhere for localised peened aluminium areas on the
aircraft. <strong>The</strong> areas peened on the FS488 bulkhead are highlighted in Figure 4. Various
publications, Graham (1986), Ward (Jan 1991, July 1991), describe the process used in
peening the original FS488 bulkhead. A contractor to Northrop performed all peening on an
as-required basis.
<strong>The</strong> procedure is outlined as follows and is based on Process Specification 14023 (PS14023).
1. <strong>The</strong> vendor shot peened the outer surface <strong>of</strong> the flange from the Z101.50 line to the lower
edge <strong>of</strong> the longeron recess radius using steel shot (PS-14023) or flapper wheel (PS-
14023.1), Figure 5. This only occurred on the pre-6 inch radius bulkheads. PS-14023 says,
“shot peen at 6-10A intensity using 230 or 280 size steel shot”. <strong>The</strong> same intensity is
required if using the flapper wheel process.
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2. <strong>The</strong> bulkheads were then IVD coated and painted.
3. <strong>The</strong> outer flange area, flange fillet radius and inner surface <strong>of</strong> the flange were then glass
bead peened down to the Z101.50 line. PS-14023 states “shot peen at 6-10A using size 4
glass beads” . Note: prior to peening the IVD and paint was to be removed. However,
there is some confusion over whether this is the case. In some cases the paint is
chemically stripped leaving the IVD before final peening and in other cases, the paint and
IVD are removed by low intensity (4A) glass bead peening before the final peening.
4. <strong>The</strong> area covered by the glass bead peening is slightly different from the original peened
area, Figure 6.
5. <strong>The</strong> process specification is PS 14023.
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Figure 4: This figure shows a schematic <strong>of</strong> the FS488 bulkhead wing attachment area. <strong>The</strong> original
flapper wheel or steel bead peened area is shown by the hash marks. From drawing
74A324205 Sh 3 Revision L.
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Figure 5: <strong>The</strong> flapper wheel peening specifications from MDA PS14023.1. This process was used on
all the early F/A-18 aircraft.
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Figure 6: <strong>The</strong> glass bead peened area from drawing 74A32405 Sh 3 Revision U is shown by the
hatching marks. <strong>The</strong> original peening area is the cross-hatched markings. This figure
when compared with figure 4 (revision L) shows the increased peened area on the FS488
bulkhead as F/A-18 production increased.
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RAAF F/A-18 aircraft A21-1 through to A21-18 and A21-101 through to A21-114 were
manufactured with a 143-pr<strong>of</strong>ile flange bulkhead. This 143-pr<strong>of</strong>ile flange was reworked to 6
inch pr<strong>of</strong>ile flange during Engineering Change Program 45 (ECP45), along with the longeron
recess area, Figure 7. <strong>The</strong> RAAF contractor, with RAAF agreement, and acting on advice
from AMRL derived from early research (Clayton and Clark 1988), used a procedure which
differed from the Original Equipment Manufacturer (OEM) specification (Anderson, 1990;
Barter, 1990 and Stanier, 1990). <strong>The</strong> OEM process calls for the rework peening to be
performed over the original vendor peening.
6"radius pr<strong>of</strong>ile
F143 pr<strong>of</strong>ile
Shaded areas indicate the F143
flange pr<strong>of</strong>ile segments removed
during the 6" radius pr<strong>of</strong>ile
modification.
Figure 7: A schematic showing the location <strong>of</strong> <strong>of</strong>fcuts from ECP45 delivered to AMRL from 32
aircraft.
In 32 RAAF F/A-18 aircraft which went through ECP 45, the whole area originally peened
by the vendor was polished before re-peening. <strong>The</strong> critical area <strong>of</strong> the bulkhead was to be
polished to a surface roughness <strong>of</strong> 0.2 micrometres (8 microinches) as defined by ASME B46-
1985. This involved an initial dry polish with 250# Aluminium oxide abrasive paper, then a
final polish with P400 (360#) Aluminium oxide abrasive paper, Figure 8. A total loss glass
bead system was also used instead <strong>of</strong> a recycled and filtered glass bead system, Figure8.
Summarising:
1. AMRL rework process: OEM peen → polish → total loss glass bead re-peen.
2. OEM rework process: OEM peen → OEM peen (peen directly over original peen)
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Figure 8:
<strong>The</strong> polished FS488 bulkhead before peening (left) showing the polished surface and the
finished product (right) after peening during ECP45.
<strong>The</strong> reason for this change <strong>of</strong> process was AMRL concern that peening over a damaged
surface would probably do nothing to improve fatigue life and might in fact degrade the
component fatigue life. <strong>The</strong> recommendation from AMRL was;
“Given that when shot peening is ineffective it is because <strong>of</strong> a poor surface finish from
the peening, and that the original machined pr<strong>of</strong>ile (143 pr<strong>of</strong>ile) will contribute to this
effect, it is recommended that the critical flange regions be polished then shot peened.
This recommendation does not guarantee any life improvement, but is made on the basis
that it is more likely to give an improvement than the alternative.”
Sharp et al. (1996), later examined in detail the OEM peened surface finish quality from FS488
bulkheads <strong>of</strong>f-cuts provided during ECP45 (machining 143 pr<strong>of</strong>ile to 6 inch pr<strong>of</strong>ile). <strong>The</strong>
results indicated that the surface finish was very poor, with lots <strong>of</strong> embedded glass
fragments, and some areas showing no indication <strong>of</strong> peening.
<strong>Peening</strong> is now a standard localised rework treatment on the F/A-18 and is used on
numerous structural components. Since the FS488 bulkhead rework process was developed,
the biggest change has been the progression to ceramic bead peening. Ceramic beads are
denser than glass and do not fracture as readily, though it is noteworthy that the peening
conditions ie. Almen intensity have not changed. Northrop has conducted research into this
area (Lambase (1990), Gallardo (1991) and Brenashan et al. (1990 and 1991)). <strong>The</strong> results
indicate an improvement in fatigue life using ceramic beads compared with glass bead
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peening, due to the reduction in surface defects which are caused by glass bead fragments
embedding into the surface. Northrop also determined from experiments that a peening
intensity <strong>of</strong> 8A provided optimum life improvement for 7050 material. Ceramic bead
peening is now used on all RAAF aircraft where localised peening is performed. To ensure
complete coverage, a layout ink is used on the area to be peened; this ink allows improved
assessment <strong>of</strong> the area coverage. <strong>The</strong> area is peened twice to ensure saturation and coverage.
<strong>The</strong> introduction <strong>of</strong> these extra variables (a change in peening media and changed intensity)
on the fleet aircraft also complicates the matter <strong>of</strong> determining a life improvement factor (ie
fatigue life as a multiple <strong>of</strong> the un-peened fatigue life).
Using a specimen that replicated the FS488 bulkhead lug area and a test surface fatigue stress
defined by k t s=69ksi 1 in the critical area the results listed in Table 1were obtained by MDA.
<strong>The</strong> specimens were peened using a shot bonded flapper wheel to an Almen intensity <strong>of</strong>
between 0.006-0.010A.
Table 1:
Results supplied by McDonnell Douglas Aircraft Company for a peened specimen
simulating the FS488 wing attachment lug region.
Peened Life (hours) Average Life (hours)
No 16960
15030 15995
Yes, initially 21460 21460
Yes, initially and every 6300 hrs 36630 36630
From these results MCAIR arrived at a single shot peening average life improvement factor
(LIF) <strong>of</strong> 1.34 and a multiple shot peening average LIF <strong>of</strong> 2.29, although very few coupons
were tested. This was then used to support fatigue life estimation for the F/A-18 aircraft.
In another series <strong>of</strong> experiments, examining a number <strong>of</strong> variables (pre-cycle [ie. fatigue
cracking before treatment] and bead type) Smith (1990) made the following observations in
relation to pre-cycling and glass and ceramic bead peening;
• For specimens pre-cycled to 2800 Simulated Flying Hours (SFH), shot peening only is
required to yield more than 12000SFH <strong>of</strong> crack initiation life (life for a crack to reach
0.01inch),
• For specimens pre-cycled to 4200SFH both shot peening and a “bonded doubler” are
required to yield more than 12000SFH <strong>of</strong> crack initiation life,
• Regardless <strong>of</strong> stress level the shot peening LIF is 3.0 (average) for these experiments.
<strong>The</strong> load spectrum used was the USN ST-16. What was particularly interesting about this
result is the large improvement in average LIF from 1.34 for flapper wheel to a value <strong>of</strong> 3.0
for glass and ceramic bead peening. It was also interesting that while numerous glass beads
fragments were found embedded in the surface, in only two cases (out <strong>of</strong> 6) did the fatigue
1 At all times during the report the specimen k t s stress will be quoted. Appendix 1 has the
conversion factors for determining applied stress.
13

DSTO-RR-0208
crack initiate from an embedded fractured glass bead. Many <strong>of</strong> the original F/A-18 aircraft
had flapper wheel peening, this was changed in the process specification to glass/ceramic
bead peening in the early eighties.
MDA claim an average LIF <strong>of</strong> 1.39 for shot peening for the F/A-18 aircraft. However,
NAVAIR, who bought the aircraft now operated by RAAF, do not allow any fatigue “credit”
for shot peening because <strong>of</strong> the high load and assumed process variability. It is this process
variability that needs to be examined in detail to fully justify the use <strong>of</strong> a specific life
improvement factor. This report will discuss the results achieved in AMRL research in the
examination <strong>of</strong> peening variability and will outline areas that still need to be researched.
3. AMRL Results
3.1 Overview <strong>of</strong> <strong>Peening</strong> Research Programs
<strong>The</strong> early stages <strong>of</strong> the program (Clayton and Clark (1988)), discussed in more detail later,
focussed on evaluation <strong>of</strong> peened and unpeened specimens which were subsequently fatigue
tested under constant-load amplitude conditions; further tests (Clark and Clayton (1991))
extended the results to spectrum loading. This research identified the significant role played
by the surface damage caused during peening <strong>of</strong> aluminium alloys, and the fact that such
damage can markedly reduce (or even reverse) the benefits provided by the sub-surface
residual stress distribution. <strong>The</strong>se early tests did not involve large numbers <strong>of</strong> replicate
testing, and so were not useful for statistical determination <strong>of</strong> LIF factors. <strong>The</strong>y did,
however, identify the critical metallurgical surface features, and led directly to the decision
to polish <strong>of</strong>f the original (OEM) peening when a rework was performed on some RAAF
aircraft, the intention being to re-apply peening without the risk <strong>of</strong> simply hammering
existing defects further into the surface.
<strong>The</strong> second part <strong>of</strong> the program involved an examination <strong>of</strong> the surface quality <strong>of</strong> the RAAF
fleet aircraft. <strong>The</strong> rework <strong>of</strong> several aircraft led to small slivers <strong>of</strong> material being available
from those bulkheads.
A third phase <strong>of</strong> the program involved attempts to find a quantitative relationship between
surface roughness condition and post-peening life, and investigation <strong>of</strong> possible correlation
<strong>of</strong> sub-surface residual stress measurements with fatigue crack growth.
<strong>The</strong> most recent research has focussed on developing the early approach <strong>of</strong> surface removal
and re-peening into a defined mid-life rework peening method, defining the LIF achievable
for RAAF F/A-18 aircraft, and on examining several critical geometrical arrangements which
could conceivably lead to anomalous results.
3.2 Early research - definition <strong>of</strong> potential problem
It is well known that steel shot bead peening is beneficial to the fatigue life <strong>of</strong> ferrous
materials, particularly the high strength steels. <strong>The</strong> benefit <strong>of</strong> steel shot bead peening <strong>of</strong>
14

DSTO-RR-0208
aluminium alloys has been assumed, but is far less well defined. Clayton and Clark (1988)
obtained results which indicate that steel shot bead peening <strong>of</strong> aluminium alloys degraded
the fatigue initiation and early growth performance, compared with a polished surface,
Figure 9. <strong>The</strong>y also observed the benefits <strong>of</strong> peening aluminium alloys with glass beads,
Figure 9, and what appeared to be a broad correlation between surface roughness and
fatigue life, although at that time the surface roughness could not be accurately quantified,
Figure 10.
Unpeened tests Glass bead peen Steel Bead peen #110
Steel Bead peen #280 Poly. (Unpeened tests ) Poly. (Glass bead peen )
Poly. (Steel Bead peen #110 )
500
450
400
Cyclic Stress (MPa)
350
300
250
200
150
100
3 3.5 4 4.5 5 5.5 6 6.5
Log Cycles (Constant amplitude)
Figure 9: Constant amplitude fatigue life results comparing, a) steel shot peened specimens with a
polished surface, b) glass bead peening under constant amplitude loading and c) a
summation <strong>of</strong> all the results showing reduced lives for steel bead peening tested compared
to glass bead peening (Clayton and Clark (1988).
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DSTO-RR-0208
Figure 10: SEM photographs <strong>of</strong> the peened surface after, Top- steel shot bead peening and Bottomglass
shot bead peening, Clayton and Clark (1988).
<strong>The</strong> results <strong>of</strong> this work highlighted the potential for surface damage from peening to
accelerate the early development <strong>of</strong> fatigue cracks; fatigue life was therefore the result <strong>of</strong> a
competition between this additional damage and the beneficial effects <strong>of</strong> the sub-surface
residual stresses.
From the early research programs, AMRL was able to determine which peening process
variable caused different types <strong>of</strong> surface damage to the specimen. A poor glass bead
filtering process caused the surface to become “chopped-up” and “pock-marked” due to
impact from broken beads, Figure 11. Microstructural grains start to lift away from the
specimen surface when an over-pressure peen or excessive-time peen is part <strong>of</strong> the process.
If the peening nozzle is not held near-vertical to the surface, smearing is caused on the
specimen surface. Specimens peened with steel shot obtain a “reddish” tinge from the iron
oxide debris that is left on the surface. Generally with glass beads there is significant debris
trapped in the surface, while the use <strong>of</strong> ceramic beads leads to only small amounts <strong>of</strong>
trapped debris.
16

DSTO-RR-0208
Figure 11: <strong>The</strong> effect <strong>of</strong> different peening variables on surface finish, Top- over-pressure or excessive
time, Bottom left- nozzle angle and Bottom right- even indentations but numerous surface
cuts from debris.
Aircraft are subject to spectrum loading and there is the potential (a) for a high spectrum
loads to change the beneficial residual stresses in the component and (b) for different fatigue
lives as a result <strong>of</strong> a changed balance between crack initiation and crack growth. Clark and
Clayton (1991) examined the effect <strong>of</strong> constant amplitude and spectrum loading on the
fatigue life <strong>of</strong> peened components, using small numbers <strong>of</strong> specimens, at different overall
stress levels. A more detailed test program examining the effect on fatigue life <strong>of</strong> peening
treatment under spectrum loading was carried out at AMRL (Sharp, Clayton and Clark,
1993). <strong>The</strong> test results, Figure 12, clearly show that at the high stress (low cycle) end <strong>of</strong> the
fatigue curve, shot peening can be detrimental to fatigue life. At the low stress (high cycle)
end <strong>of</strong> the fatigue curve, shot peening is beneficial to fatigue life but the amount <strong>of</strong> benefit is
controlled by the peening process quality.
17

DSTO-RR-0208
550
USN Spectrum (ST-16)
Peak Spectrum Stress (MPa)
500
450
400
350
Intersection <strong>of</strong> contractor peen
specimens and polished specimens
800# Hand Polish
Contractor Peen
AMRL Laboratory Peen
AMRL OEM Simulation Peen
Intersection <strong>of</strong> AMRL laboratory peen
specimens and polished specimens
300
1 10 100 1000
Fatigue Life (Blocks)
Figure 12: Test results comparing a variety <strong>of</strong> peening processes under spectrum loading (Sharp,
Clayton and Clark, 1993). <strong>The</strong> figure illustrates the potential for peening to reduce the
fatigue life at all stresses. <strong>The</strong> spectrum used was ST-16, a US Navy spectrum developed
for the F/A-18.
<strong>The</strong> F/A-18 aircraft was not designed to fly with this spectrum at the high stresses
(>450MPa) and is lifed on the 350MPa to 420MPa range. A possible fatigue life problem
arises due to the “cross-overs” observed in Figure 11; what is the extent <strong>of</strong> any benefit? <strong>The</strong>
AMRL peen (laboratory peening) is very good and provides significant improvement in
fatigue life compared to a good polish (800#) over the 350-450MPa stress range. <strong>The</strong>
commercial peen (peening by a commercial operator) is no better than a good polish above
410MPa and but shows an improvement <strong>of</strong> the original peening simulation below 460MPa.
<strong>The</strong> AMRL abusive peen (an attempt to simulate original OEM peening) would never be
better than a good polish over the complete range <strong>of</strong> testing stresses. Clearly, the benefits
from peening are linked to the applied stress; at high stresses there is only a slight
improvement (if any) whereas at low stresses there can be a significant improvement in
fatigue life. This accords with normal fatigue observations in that, at low stresses, fatigue life
is particularly sensitive to surface condition (i.e. the “lead” crack nature defines the life), and
in peening, the crack retardation process caused by the surface residual stresses can play a
major role. At higher stresses, however, many cracks initiate rapidly, and life is controlled
by crack growth.
<strong>The</strong> specimens were examined metallographically to determine the fatigue initiation point,
Figure 13, and the surface roughness (see later).
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DSTO-RR-0208
Figure 13: Typical fatigue crack initiating defects from different peening processes. <strong>The</strong> majority <strong>of</strong>
fatigue cracks initiate at embedded fractured glass beads, surface laps and folds and
fractured sub-surface intermetallics.
Experiments conducted by a number <strong>of</strong> researchers (Butz and Lyst 1961, Fuchs 1956 and
Almen 1947) have suggested the use <strong>of</strong> “dimple diameters” to estimate depth <strong>of</strong> peening
effect. <strong>The</strong> results show essentially a 1-to-1 relationship between dimple diameter and depth
<strong>of</strong> compressed layer, with the residual stresses dropping away rapidly after this point. In
more recent research (Al-Hassani 1984) it has been shown that in high strength steels this 1-
to-1 ratio applies but in aluminium alloys ratios <strong>of</strong> 1-to-2 have been obtained.
A research program was also conducted at AMRL to obtain a better understanding <strong>of</strong>
peening parameters and their effect on fatigue life. In particular gas pressure, irradiation
time, bead flux and bead type were examined. <strong>The</strong> tests were conducted on small 4-point
bend specimens (100mm long, 5mm wide and 3mm thick). Each specimen was polished to
#800 grade finish before treating. All testing was done at 10Hz and R=0.05 and a test stress
<strong>of</strong> 186MPa. Seven treatments, Table 2, were compared to a polished finish.
19

DSTO-RR-0208
Table 2: Four point bend tests to determine effect <strong>of</strong> a range a peening parameters quickly.
Treatment No. Gas Pressure Irradiation Bead Flux Bead Type
(Psi) Time (seconds) (m 2 /second)
1 74 40 1.2x10 7 Used glass
2 74 40 1.2x10 7 Fresh glass
3 74 19 1.2x10 7 Fresh glass
4 74 58 1.2x10 7 Fresh glass
5 95 40 1.2x10 7 Fresh glass
6 50 40 1.2x10 7 Fresh glass
7 30 40 1.2x10 7 Fresh glass
<strong>The</strong> results <strong>of</strong> these experiments are shown in figure 14.
Unpeened Treatment 1 Treatment2 Treatment3 Treatment4 Treatment5 Treatment6 Treatment7
7
6
5
4
Treatment
3
2
1
Unpeened
0 20000 40000 60000 80000 100000 120000 140000 160000
Fatigue life (Cycles)
Figure 14: Comparisons <strong>of</strong> hand-polish with various peening treatments looking at the effects <strong>of</strong>
peening process variables.
Treatment 2 is the standard treatment proposed for the RAAF aircraft, there is clearly a
significant life improvement over polished specimens and over used glass beads (treatment
1). <strong>The</strong> fatigue life improvement is the same for treatment 2 and 3, despite treatment 3
having 100% coverage compared to 200% coverage. An increase in coverage (300%)
treatment 4 shows a reduction in life compared to treatment 2 and 3. This is the “over-peen”
case. Gas pressure seems to have little effect on mean fatigue life except in terms <strong>of</strong>
increasing the fatigue life scatter for the lower pressures.
<strong>The</strong>se simple tests showed that the peening parameters DSTO needed to examine in detail to
provide a LIF for the RAAF fleet were Almen Intensity and coverage.
20

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3.2.1 Assessment <strong>of</strong> RAAF aircraft
Twenty-four <strong>of</strong>f-cuts were examined in a SEM for surface roughness (Sharp et al. (1996 &
1997)) and a selection is shown in Figure 15a. Comparison <strong>of</strong> the original peen surface
quality with a replica <strong>of</strong> the RAAF rework surface quality, Figure 15b, clearly reveals a
dramatic reduction in cuts and other surface damage, and well-formed peening dents in the
rework surface.
A further observation was the wide range <strong>of</strong> peened surface conditions observed on the
service aircraft, indicating a high level <strong>of</strong> process variability, Figure 15.
Aircraft A21-101 Aircraft A21-103
Figure 15a: Scanning electron microscope images <strong>of</strong> a typical RAAF F/A-18 fleet surface finish after
OEM peening, (Sharp et al., 1997). Figure 15b overleaf.
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DSTO-RR-0208
Figure 15b: Scanning electron microscope pictures <strong>of</strong> a replica from a RAAF reworked surface,
Aircraft A21-110 (left) and Aircraft A21-15 (right) (Sharp et al. 1997).
3.2.2 Surface roughness/residual stress measurement
Since a smoother surface appeared to provide substantial improvements in fatigue life, it was
desirable to explore the possibility that a correlation could be found between surface
roughness condition and fatigue life.
AMRL purchased a laser surface pr<strong>of</strong>iler, which allowed very accurate non-contact
measurements <strong>of</strong> the surface roughness. <strong>The</strong> laser surface pr<strong>of</strong>iler is accurate to 0.1µm over
a range <strong>of</strong> 500µm. While this level <strong>of</strong> accuracy is not required in this case, many “stylus” type
<strong>of</strong> surface roughness measuring systems grind through the surface due to the complex
surface roughness. <strong>The</strong> AMRL instrument is now computerised to allow a 3-D surface
pr<strong>of</strong>ile to be built up.
For the RAAF aircraft <strong>of</strong>f-cuts, the wide difference in surface quality is very evident in the
measured surface roughness, Table 3, where R max is the max roughness depth, R z is the mean
roughness depth and R a is the arithmetical mean deviation (See Appendix 1).
Table 3: Some Typical F/A-18 FS488 Bulkhead Off-cut Surface Roughness Measurements (Sharp et al
1997).
Tail No. R max (µm) R z (µm) R a (µm)
A21-13 63.96 30.95 3.66
A21-18 73.12 41.36 5.03
A21-109 103.1 35.43 3.72
A21-112 124.6 55.53 6.23
Table 4 compares coupon surface roughness with the coupon fatigue life at 410MPa (F/A-18
peak applied stress under a USN ST-16 load spectrum, which gives a k t s=432MPa).
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Table 4: Comparison <strong>of</strong> fatigue life and surface roughness for a number <strong>of</strong> peening processes. <strong>The</strong>
specimens were tested using the ST-16 load spectrum with a peak stress <strong>of</strong> 432MPa.
Surface Process R max (µm) R z (µm) X-ray Residual
Stress (MPa)
Fatigue
Life
(Programs)
AMRL Glass Bead 19.7 14.5 -230±30 150
Peen
Commercial Glass 28.2 21.5 -230±30 65
Bead Peen
AMRL Steel Shot 50.3 40.5 -230±30 55
Peen
AMRL Polish #800 3.4 2.9 -20±30 63
<strong>The</strong> results appear to show a direct correlation between surface roughness and fatigue life <strong>of</strong>
peened aluminium alloys; as roughness increases, life decreases,Figure16. It was therefore
considered that it might be possible to quantify the effects <strong>of</strong> F/A-18 fleet peening on fatigue
life by measuring surface roughness from replicas obtained from peened areas. However,
further results (discussed below) obtained using a more aggressive peening method
indicated that the apparent relationship broke down, the problem being that excessive or
aggressive peening leads to defects being buried sub-surface and to very deep laps, which do
not influence the surface roughness measurement, while greatly reducing fatigue life. <strong>The</strong>
conclusion at this stage is that some correlation exists between roughness and life, but where
the history <strong>of</strong> the part may have included some aggressive peening, this relationship is
unreliable.
Roughness vs Fatigue Life
Roughness (micrometre)
100
80
60
40
20
0
Roughness Rmax
Roughness Rz
0 50 100 150 200
Fatigue Life (Blocks)
RAAF
Aircraft
Figure 16: <strong>The</strong> relationship between measured surface roughness and fatigue life for three different
peening conditions. <strong>The</strong> surface roughness on RAAF aircraft samples was R max =64-125µm
and R z =30-60µm as marked, indicating a possible shorter fatigue life than the laboratory
conditions indicated.
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DSTO-RR-0208
Research by Fair et al., (1984), indicated that service loading <strong>of</strong> a component can diminish or
completely remove the residual compressive stresses thus reducing any fatigue life benefits.
This reduction is particularly likely to happen if the component is subjected to large
compressive loads. However, work by Hammond and Meguid, (1990) on 7075-T7351 showed
no relaxation in residual stress near the surface as a function <strong>of</strong> used life. Oshida and Daly
(1990), in tests on 7050-T74651 showed a distinct relaxation in both surface and interior stress
with fatigue life. In both research programs, the peened specimens showed an improved
fatigue life over a range <strong>of</strong> stresses, they both used 4-point rotating bending specimens and
yet they achieved different results. Figure 17, is typical <strong>of</strong> the change in residual stress with
depth due to spectrum loading observed by a number <strong>of</strong> researchers. <strong>The</strong> exact cause <strong>of</strong> this
“shake down” effect has yet to be determined, mainly due to the large variation in the
results.
300
O flights 800 flights 10000 flights
200
Residual Stress (MPa)
100
0
0 0.5 1 1.5 2 2.5
-100
-200
-300
-400
Residual Stress Depth (mm)
Figure 17: X-ray residual stress result, showing the effect <strong>of</strong> spectrum loading on the residual
compressive stress caused by peening an aluminium alloy (Fair et. al, 1984). <strong>The</strong>re
appears to be a distinct shakedown effect as the number <strong>of</strong> flights increases.
Clark and Clayton (1991) also undertook residual stress analysis <strong>of</strong> their peened specimens.
<strong>The</strong> results are shown in Figure 18. Figure 18a shows the x-ray residual stress pr<strong>of</strong>ile <strong>of</strong> a
peened specimen. Figure 18b and Figure 18c show the effect on the residual stress pr<strong>of</strong>ile <strong>of</strong>
fatigue loading. <strong>The</strong> variability <strong>of</strong> the residual stress measurements adds to the uncertainty
as to whether the apparent reduction at greater depths was real, although the change in
surface stress appeared to be significant. <strong>The</strong> effect <strong>of</strong> loading on reduction <strong>of</strong> residual
stresses in aluminium alloys is still an area requiring further research.
A detailed analysis <strong>of</strong> peened surface residual stress measurements methods was made at
AMRL using a number <strong>of</strong> different techniques, ie. x-ray diffraction, neutron diffraction, hole
24

DSTO-RR-0208
drilling etc (Olsson-Jacques 1999). A recent paper by (Nobre, Kornmeier, Dias and Scholtes,
2000) also compares incremental hole-drilling (IHD) with XRD. <strong>The</strong>y indicate that IHD can
be correctly used to measure the residual stress <strong>of</strong> peened components, but supported the
AMRL conclusion that care must be taken due to the strong stress gradients.
50
No applied stress 400MPa/10000cycles 315MPa/40000cycles
0
0 0.2 0.4 0.6 0.8 1 1.2
Residual Stress (MPa)
-50
-100
-150
-200
-250
-300
Depth (mm)
No applied stress 400MPa Peak Stress 475MPa Peak Stress
50
0
0 0.2 0.4 0.6 0.8 1 1.2
Residual Stress (MPa)
-50
-100
-150
-200
-250
-300
Depth (mm)
Figure 18: Residual stress (x-ray diffraction) pr<strong>of</strong>ile for peened aluminium alloy, Clark and Clayton
(1991) –top) – non-loaded pr<strong>of</strong>ile compared with constant amplitude pr<strong>of</strong>ile, 10000cycles
at 400MPa and 40000 cycles at 315MPa and bottom) non-loaded pr<strong>of</strong>ile compared with
variable amplitude loaded specimen pr<strong>of</strong>ile, 30 blocks at 475MPa and 30 blocks at 400MPa
peak stress.
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3.2.3 Summary <strong>of</strong> AMRL <strong>Peening</strong> Specification
AMRL specified a repair procedure, which was used on approximately 32 aircraft during
ECP45. <strong>The</strong> AMRL recommendation was that any future rework peening be done to
PS14023 with the following changes;
1) ALMEN intensity 6-8A,
2) 200% coverage,
3) Nozzle perpendicular to the surface,
4) Total loss glass bead system or recycled (filtered) ceramic beads,
5) Start peening in corners and work to flat region to reduce laps and folds.
This new specification is a refinement <strong>of</strong> the original MDA PS14023 and was published in
early 1994 (Sharp et.al, 1994). At the time it was also considered important to remove any
previous peening “damaged material” from the surface, though no experiments at that time
had been undertaken to confirm this approach. This aspect has been pursued to provide a
more broadly applicable rework method, and has been focussed on a repair for the FS470.5
X19 region <strong>of</strong> the F/A-18 aircraft (Sharp et. al, 2000). This particular application has been
discussed in detail elsewhere.
<strong>The</strong> development <strong>of</strong> a full rework approach, including surface removal, is described in the
following section.
3.3 Rework <strong>Peening</strong> Research
<strong>The</strong> use <strong>of</strong> a surface removal process prior to re-peening a reworked area on some RAAF
aircraft prompted research aimed at determining whether this method could be developed
into a general procedure for reworking local areas <strong>of</strong> an airframe part-way through life to
allow the fatigue life <strong>of</strong> the aircraft to be maximised. Such a method would have to be
reliable and sufficiently risk-free to allow application to critical airframe components.
Reworks on real aircraft fleets are a competition between “cost” and “benefit” in terms <strong>of</strong> life
improvement. AMRL has always stated that any damaged layer must be removed to
maximise the LIF after rework no matter whether the surface is left polished or peened.
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DSTO-RR-0208
50
Total Fatigue Life, Programs
40
30
20
10
0
0 20 40 60 80 100
Percentage <strong>of</strong> Fatigue Life Expended before <strong>Peening</strong>, %
Figure 19: <strong>The</strong> fatigue life <strong>of</strong> specimens peened after various periods <strong>of</strong> service. <strong>The</strong> specimens were
fatigue tested to a percentage <strong>of</strong> total life then peened before continuing the test until
failure. (Note: there was no surface removal between fatigue testing and peening).
Figure 19 clearly shows the benefit in removing the damage layer to restore maximum life.
<strong>The</strong>re is a significant drop in LIF if the damaged layer is not removed after the component
had reached 40-50% <strong>of</strong> its full life. In fact, the peening has almost no LIF effect, because the
damage is now deeper than the beneficial compressive residual stresses. Section 3.3 shows
that even greater care must be taken to remove a peened damage layer, otherwise not only
can no LIF can be associated with the repair process but in fact a reduced life may occur at
some stresses.
3.3.1 Different rework process comparisons
In many localised regions, the F/A-18 has been peened or re-peened, usually by contractors.
AMRL therefore examined the peening rework processes performed on the F/A-18 by these
contractors. Due to the repair requirements for the full-scale fatigue test articles in the
International Follow On Structural Test Program (IFOSTP) run by the RAAF and Canadian
Forces (CF) several <strong>of</strong> rework processes had to be examined. AMRL recommendation was
that the OEM original peening be removed by polishing before any re-peen, and that a total
loss bead system be used; in contrast the CF had adopted the OEM rework procedure <strong>of</strong>
peening directly over the original peening without polishing. In view <strong>of</strong> the potential for this
approach to bury damage in the surface, and cause excessive peening damage, this process
clearly had to be one <strong>of</strong> those examined. Sharp, Clayton and Clark (1994), compared the
effects <strong>of</strong> the different rework processes on fatigue life (Figure 20, summarised in Table 5).
27

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Peak Spectrum Stress (MPa)
420
415
410
405
400
395
390
385
Comparison <strong>of</strong> Fatigue Lives between
Ceramic Bead <strong>Peening</strong> and Glass Bead <strong>Peening</strong>.
Hand Polish
Aircraft glass bead peen simulation
Rework peen directly over original peen
Rework peen over polished surface
Ceramic bead rework peen over polished surface
380
6000 8000 10000 30000
Fatigue Life (Simulated Flight Hours)
Figure 20: Experimental results comparing the fatigue lives <strong>of</strong> a number <strong>of</strong> different localised rework
peening processes. <strong>The</strong> ceramic beads where manufactured from aluminium oxide. <strong>The</strong>
beads were very spherical, had a small variation in diameter compared to glass and very
hard and heavy.
Table 5 also displays the surface roughness results and the residual stress results for each <strong>of</strong>
the different peening rework processes. <strong>The</strong> residual compressive stress is a factor in the
improved fatigue life due to peening, and AMRL and Monash University (Victoria), built a
rig which could hold the test specimens for x-ray residual stress testing. This rig simplified
the process immensely, meaning that residual stress results could be obtained easily and
with improved accuracy.
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Table 5:
Results from local peening rework processes. <strong>The</strong> OEM rework – peen over OEM original
peen, AMRL rework – remove OEM original peen by polishing then peen. Specimens
(Sharp, Clayton and Clark (1994)) were tested with a RAAF post-LEX spectrum Spec 16 -
21010 turning points 302.9 hrs per block at a peak stress 410 MPa. Material was 5 inch
plate,
Rework Process Log Av. Fatigue Surface Roughness Surface Roughness Residual Stress
Life (SFH)
R max (µm)
R z (µm)
MPa
Hand polish (#800) 9287 3.4 2.9 -35±20
Simulated original 19079 63.9 46.5 -270±40
aircraft peen
“Overpeen” rework 15073 51.7 38.0 -230±40
AMRL rework
specification
22895 28.2 21.4 -260±40
Table 5 shows that the residual compressive stress is very similar no matter which treatment
is used - in other words they all have reached saturation. Interestingly, in all the x-ray
residual stress analyses completed, the residual compressive stress range is between -
230MPa and -270MPa. <strong>The</strong> ranking order is very similar to the order presented in Table 4,
which were fatigue tests, conducted using a different spectrum ST-16.
Figure 21 shows typical surface finishes achieved with each <strong>of</strong> the peening rework processes
trialled. <strong>The</strong> AMRL rework process (polish/peen) provides the best surface finish and also
the longest fatigue life.
Figure 21a: A typical OEM peen simulation surface. Note that because glass beads were used in the
simulation, the damage to the surface could not be made to fully match the damage seen on
some aircraft where a flapper wheel (Figure 50 was used. <strong>The</strong> arrow indicates lifting grains
typical <strong>of</strong> over-pressure or excessive peening time.
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DSTO-RR-0208
Figure 21b: A typical surface finish, resulting from a re-peening over the OEM original peening. <strong>The</strong>
arrow indicates “chopped” up surface due to glass bead fragments impacting the surface.
Figure 21c: A typical surface finish, resulting from removing the original OEM peening and then repeening.
<strong>The</strong> arrow indicates a small gap in the surface due to the peening process.
Despite this the surface has regular even indentations seen on a good peened surface.
<strong>The</strong> newer results indicated that there are some conditions where surface roughness could
not be directly correlated to fatigue life; the OEM specified rework (peening directly over the
original peened surface) was a particular problem in that the over-peening reduces the
surface roughness but also reduces fatigue life. <strong>The</strong> reason for this result is clearly visible in
Figure 22; the original peened surface (Figure 22a), is very rough compared to the AMRL
rework polish/peen (Figure 22c), and this can be related to their fatigue lives. However, in
the over-peening rework case, Figure 22b, the original surface defects are smeared into the
surface and become sub-surface defects, which cannot be measured by surface roughness.
30

DSTO-RR-0208
Figure 22a: A typical surface pr<strong>of</strong>ile from the OEM peen simulation.
Figure 22b: A typical surface pr<strong>of</strong>ile from re-peening directly over the original OEM peen.
Figure 22c: A typical surface pr<strong>of</strong>ile from the AMRL rework. <strong>The</strong> original OEM peening damage has
been removed and the surface re-peened.
This observation indicates that if the component has been peened directly over the original
peening then surface roughness cannot be correlated with fatigue life. This non-correlation
causes a major problem because, without a very confident knowledge about the peening
history and conditions, the peening quality and fatigue life can no longer be simply
quantified from the surface roughness.
To show more directly the effect <strong>of</strong> peening on fatigue crack growth rate, fractographic
analysis <strong>of</strong> the striation marks left on the fracture surface was undertaken. <strong>The</strong> results are
shown in Figure 23 and Figure 24.
It can be seen clearly that in the two peened specimens the initial crack growth rate is much
lower while the crack is contained in the first 200-300µm. After this depth, the crack growth
rate <strong>of</strong> the peened specimens increases, approaching that <strong>of</strong> the unpeened specimens,
becoming very similar by a crack depth <strong>of</strong> 600µm. Note that the specimens were all tested at
a peak applied stress <strong>of</strong> 410MPa and not at the high stresses (>450MPa) shown earlier where
peening had the potential to reduce specimen fatigue life compared to polished specimens.
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DSTO-RR-0208
10
Crack Depth (mm)
1
0.1
Polished
Peened
0.01
0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85
Program Number
Figure 23: Measurement <strong>of</strong> fatigue crack depth (mm) versus program number. <strong>The</strong> figure indicates
that the initial defect is larger in the peened specimens, but that the crack growth
increment per program is very low in the early stages, illustrating the two competing
processes which determine the amount <strong>of</strong> fatigue life extension.
1
Reduction in growth rate
Crack Growth Rate, da/dN
0.1
0.01
1E-3
Peened specimens
1E-4
0.01 0.1 1
Crack Length, a (mm)
Figure 24: Measurements <strong>of</strong> the crack growth rate (da/dn) versus crack length showing clearly the
effect peening has on retarding the fatigue crack growth (Sharp and Byrnes (1996).
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DSTO-RR-0208
A further important point from these results is that the crack growth rate is always lower (or
similar) in the peened specimens than in the control (non-peened) specimen i.e. the “tension”
residual stresses deeper below the surface do not (as popularly believed) accelerate crack
growth rates significantly.
3.3.2 F/A-18 fleet rework simulation
<strong>The</strong> next stage <strong>of</strong> examining the effect <strong>of</strong> the rework peening process on fatigue life was to
simulate the fleet condition in more detail, particularly the fact that many RAAF aircraft
were subjected to a rework program part-way through life. This comparison was done by
cycling the peened specimen for some period before reworking the specimen and then
cycling to failure. Table 6 lists the different fleet simulations tested in this part <strong>of</strong> the
program. <strong>The</strong> RAAF loading spectrum “Spec 16” was applied to the specimens, with a peak
stress <strong>of</strong> 410MPa. <strong>The</strong> specimens were similar to those used throughout the complete testing
program.
Table 6: Fleet simulations tested in the program.
Simulation Label
C2
F1
F2
F3
F4
Simulation Specification
AMRL Simulation <strong>of</strong> original aircraft (OEM) peen
AMRL Sim OEM peen → fatigue for 3000SFH → glass bead rework → test to failure
AMRL Sim OEM peen → fatigue for 3000SFH → glass bead rework → fatigue for
3000SFH → glass bead rework → test to failure
AMRL Sim OEM peen → fatigue for 6000SFH → glass bead rework → test to failure
AMRL Sim OEM peen → fatigue for 3000SFH → glass bead rework → fatigue for
3000SFH → ceramic bead rework → test to failure
Simulation C2 represents the AMRL laboratory simulation <strong>of</strong> the original OEM aircraft peen.
Simulation F1 represents the 32 RAAF aircraft that went through ECP45 (a bulkhead radius
rework) after some period <strong>of</strong> flying (the lead RAAF aircraft went through ECP45 at
approximately 1500 hours). <strong>The</strong> original OEM peen was removed after 3000SFH and the
specimen peened with a total loss glass bead system. Simulation F2 represents the case
where areas from simulation F1 might need to be reworked again during the life <strong>of</strong> the
aircraft. Simulation F3 represents the other 42 aircraft that did not go through ECP45 and
might need to be reworked some time during the aircraft life. Simulation F4 represents the
latest F/A-18 peening process, which uses ceramic beads instead <strong>of</strong> glass beads. <strong>The</strong> fatigue
life results are listed in Table 7.
Based on earlier crack growth work (Barter, July 1990), 200-300µm <strong>of</strong> material had to be
removed from each face during a polish rework, to ensure that no cracks would be present
before re-peening <strong>of</strong> the surface. <strong>The</strong> first column in Table 6 is the rework process, the
second column is the fatigue life from the last rework until failure, and the third column is
the total specimen fatigue life.
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Table 7: <strong>The</strong> average log fatigue life results for each peening/rework simulation.
Simulation Process
Av. Fatigue life after
last rework (SFH)
Addition <strong>of</strong> fatigue life
between reworks (SFH)
Total Av. Fatigue Life
(SFH)
C1 9287 0 9287
C2 16550 0 16550
F1 16026 3000 19026
F2 17624 3000+3000 23624
F3 20327 6000 26327
F4 17865 3000+3000 23865
<strong>The</strong> results are also shown in Figure 25. For comparison a polished specimen (C1) fatigue
life is 9287SFH. Column 2 in Table 7 shows that all peening gives an improvement in
subsequent fatigue life at the test stress k t s= 410MPa. All other tests were completed on new
coupons, which is why C2 (OEM simulated original peening) average fatigue life is slightly
lower than the results presented in Table 5 for the same condition.
30000
Polish
Peen
Fatigue 3000hrs-rework-fatigue
Fatigue 3000hrs-rework-fatigue 3000hrs-rework-fatigue
Fatigue 6000hrs-rework-fatigue
Fatigue 3000hrs-rework-fatigue 3000hrs-ceramic rework-fatigue
Number <strong>of</strong> Cycles
25000
20000
15000
10000
5000
0
Cycles to failure from last rework
Total Cycles (rework+failure)
Figure 25: Left –Average number <strong>of</strong> cycles to failure after the last rework , on the Right-Total number
<strong>of</strong> cycles to failure. As can be seen there is little variation in the cycles to failure after the
last rework, indicating that sufficient material has been removed before final peening to
restore complete life.
A significant result in terms <strong>of</strong> aircraft lifing is the reduction in fatigue life scatter when
using ceramic beads compared to glass beads. This reduction can be attributed directly to the
quality <strong>of</strong> the beads and the reduction in the number <strong>of</strong> broken beads embedding in the
specimen surface. It is also interesting that ceramic beads provide a similar fatigue life to
glass beads. This result was observed by MCAIR but not by earlier AMRL testing that
34

DSTO-RR-0208
showed a significant improvement in fatigue life using ceramic beads compared to glass
beads. <strong>The</strong> reason for the different results is not clear, but may be due the different testing
conditions or process details. Figure 25 indicates that an operator can rework a local area a
number <strong>of</strong> times and each time the original peened fatigue life is restored to that area. On
an aircraft <strong>of</strong> course, the reduction in thickness (increased stress) with each polish would
have to be taken into account.
3.3.3 Summary <strong>of</strong> AMRL Rework <strong>Peening</strong> Specification
AMRL recommendation was that any future rework peening be done to PS14023 with the
following changes;
1) Remove original peening surface damage (or other damage) before peening,
2) Polish surface prior to peening,
3) ALMEN intensity 6-8A,
4) 200% coverage in main area,
5) Nozzle perpendicular to the surface,
6) Total loss glass bead system or recycled (filtered) ceramic beads,
7) Start peening in corners and work to flat region to reduce laps and folds,
8) Ensure peening runout region ie 100% coverage.
<strong>The</strong> improvement over the specification set out in section 3.2.3 is the verification that the
removal <strong>of</strong> “damage layer” is critical to the process and the life improvement factor derived
from any peening process.
This complete process was published and distributed in late 1996 (Sharp and Byrnes, 1996)
and has been used on both fleet aircraft and structural test articles (Section 3.5). CF adopted
this process for any peening repairs to the FT55 test article. To enhance control <strong>of</strong> the
amount <strong>of</strong> material to be removed AMRL developed a spring punch process for controlled
surface removal (Barter and Houston, 1997). <strong>The</strong> process essentially places a large dent into
the surface <strong>of</strong> a controlled depth. Once the dents have been polished away a controlled
depth <strong>of</strong> material has been removed.
3.4 <strong>Peening</strong> Life Improvement Factor (LIF)
<strong>The</strong> basis <strong>of</strong> any repair is the LIF that can be derived from that repair. For the F/A-18 while
many components are peened no LIF is taken account <strong>of</strong> in the lifing. <strong>The</strong> main reason for
this has been the unreliability with the peening process. Figure 26 shows how the coupon
fatigue life varies with the peening parameters ALMEN intensity and peen coverage. <strong>The</strong>
variability in life is an indication <strong>of</strong> why robotic or automated peening is far better than
human operator peening.
Figure 26 also shows that using 6-8A ALMEN intensity at 200% coverage provides
maximum LIF for 7050-T7451 aluminium alloy coupons.
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200
180
160
8A ALMEN
6A ALMEN
No peening
Fatigue life-programs
140
120
100
80
60
40
20
Mil Spec
0
-100 0 100 200 300 400 500
Coverage (%)
Figure 26: <strong>Effect</strong> <strong>of</strong> coverage and ALMEN intensity on fatigue life using RAAF Spec16 spectrum at
an applied load <strong>of</strong> 394MPa.
More recent tests have shown a relationship between mean LIF and maximum applied
spectrum stress, as shown in Figure 27.
7
Life Improvement Factor
6
5
4
3
2
1
LIF=6*10 15 x -5.8749 R 2 =0.9791
0
300 350 400 450 500 550
Max applied Stress (MPa)
Figure 27: Results from Y470.5 X19 pocket test program using the IARP03a+marker spectrum at a
range <strong>of</strong> stresses (Sharp et. al, 2000).
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DSTO-RR-0208
Figure 27 shows a rapidly increasing mean LIF with lower stress, but detailed results
showed the lower bound LIF value increases a lot more slowly. At the stresses (400-420MPa)
examined in many <strong>of</strong> the research programs a LIF=1.5 is conservative and appropriate. A
more detailed explanation <strong>of</strong> the work done on the Y470.5 X-19 pocket can be found in
(Sharp et. al, 2000).
3.5 Fleet Applications Research
<strong>The</strong> peening process specified in Section 3.2 and Section 3.3 has been the basis for a number
<strong>of</strong> experimental test programs directly related to repairs on the F/A-18 aircraft fleet.
3.5.1 Y470 Crotch Repair Program
<strong>The</strong> Y470 crotch region is a highly stressed region <strong>of</strong> the wing carry through bulkheads.
MDA released ECP365 to repair this area and the X19 region. <strong>The</strong> repair called for metallic
doublers on the crotch region. Based on previous research DSTO devised a composite patch
repair to replace the metallic patch. Due to the severe curvature <strong>of</strong> the region the patch
design alone was a very complex process. However to maximise the benefits <strong>of</strong> the
composite patch a high temperature curing adhesive is needed. <strong>The</strong> MDA <strong>Peening</strong>
specification PS14023 RevH indicates that the temperature <strong>of</strong> peened aluminium areas
should not exceed 250 0 F (121 0 C) but does not state any time period. <strong>The</strong> objective <strong>of</strong> the test
program is to see what effect time and temperature (typical <strong>of</strong> patch application) would have
on the underlying residual compressive stress due to peening.
Interestingly the Boeing process specification BAC 5730 (1988) for peening specifies a
maximum temperature for peened aluminium regions <strong>of</strong> 200 0 F (93 0 C) above which the area
has to be repeened.
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DSTO-RR-0208
Figure 28: A schematic showing the location <strong>of</strong> the Y470 bulkhead crotch region.
<strong>The</strong> following steps were followed for to the test coupons to simulate the repair process
surface treatment. Step 7 simulates the actual heating <strong>of</strong> the adhesive FM73.
a) Ceramic bead peen to MDA specifications PS14023 RevH
b) solvent degrease to waterbreak standard,
c) mild scotchbrite abrade,
d) alumina grit-blast
4) 1% aqueous silane for 15 minutes (A187),
5) dry at 110 0 C for 1 hour,
6) BR127 primer applied, dried in air for 30 minutes and at 110 0 C for 30 minutes,
7) cure 120 0 C for 1 hour to simulate FM 73 adhesive cure.
<strong>The</strong> results <strong>of</strong> the coupon testing are shown in Table 8 and Figure 29.
38

DSTO-RR-0208
Table 8: Coupon fatigue life results. Five coupons were tested per condition.
Base Line Peen Peen + Patch Repair
Process
Coupon 1 60.85 80.19
Coupon 2 71.19 70.54
Coupon 3 108.54 55.02
Coupon 4 62.58 58.98
Coupon 5 63.58 66.87
Log Average 71.1 66.1
A statistical comparison was made <strong>of</strong> the two process sets, which showed that there was a
97.5% probability <strong>of</strong> them being from a single population. To further confirm this result,
surface residual stress analysis was performed on the coupons. Table 9 compares the base
line residual stress with the residual stress from just the surface treatment (ie missing step 7
adhesive cure) and the residual stress for the complete process.
Fatigue Life (programs)
120
100
80
60
40
20
0
max
min
BL
RRP
Figure 29: Column graph comparing the base line test results with the repair process. Note how the
range <strong>of</strong> repaired process coupons falls in the range <strong>of</strong> baseline coupon failures
39

DSTO-RR-0208
Table 9: Residual Stress Results from X-ray Diffraction Technique.
Vertical
Direction
Base Line Peen
+ Surface
Treatment
+Surface
treatment+Heati
ng Treatment
-231 MPa -220 MPa -187 MPa
Horizontal -260 MPa -254 MPa -182 MPa
Direction
Average -246 MPa -237 MPa -185 MPa
Note: Specimens heated for 1 hour at 110 0 C, cooled to room temperature and then 2 hours at
120 0 C.
A slight drop was observed in mean surface residual stress readings. However, since the
fatigue are apparently not affected it is unlikely that a significant change in the residual
stress depth pr<strong>of</strong>ile has occurred. A recent test program conducted by QETE in Canada
(Roth and Yanishevsky 1999) supports these results.
3.5.2 Fleet Y470.5 X19 Pocket Repair program
AMRL proposed to the RAAF a fleet lifing policy for the X19 region based on its peening
specification and research into surface finish. <strong>The</strong> Y470.5 X19 pocket is a highly stressed
region (approx 420MPa) <strong>of</strong> the wing-carry through bulkheads. <strong>The</strong> X19 pocket failed during
the US Navy fatigue test ST16 and is regarded as a possible life limiting area in terms <strong>of</strong>
extending the operational life <strong>of</strong> the F/A-18.
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Figure 30: Location <strong>of</strong> the X-19 area <strong>of</strong> the Y470 bulkhead wheel well flange.
Failure
Fatigue initiation sites in
web fillet tangency and
under flange.
X19 from
centreline
Figure 31: Failure location at the X19 stiffener on ST16 after approximately 12000 SFH cycling.
Failure was LH side. Looking forward.
Based on the ST16 fatigue test and FE analysis MDA proposed ECP365 to improve the
fatigue life <strong>of</strong> the X19 and the crotch area <strong>of</strong> the Y470.5 bulkhead. <strong>The</strong> US Navy released the
X19 repair as AFC No. 174. Based on this AMRL proposed a modified repair for the X19
pocket and a subsequent RAAF fleet lifing policy. <strong>The</strong> CF have completed ECP365
modification process.
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Like all the structural aluminium parts the Y470.5 bulkhead has been IVD coated for
corrosion protection. As part <strong>of</strong> this IVD production process the fresh aluminium surface is
etched. This etching has a detrimental effect on the fatigue life <strong>of</strong> coupons made from a
similar material 7050-T7451 (Barter 1999). A comparison <strong>of</strong> peened fatigue life versus etch
fatigue life, Figure 32, shows the benefit <strong>of</strong> peening compared to an etched surface over a
large range <strong>of</strong> stresses.
Etched Fatigue Life
Peened Fatigue Life
500
450
Max Stress (MPa)
400
350
300
250
0 50 100 150 200 250 300 350 400
IARP03a programs
Figure 32: Fatigue test results using IARP03a+marker spectrum comparing an etched surface finish
with a peened surface finish. <strong>The</strong> peening was done as per AMRL standards.
All testing was done under spectrum loading using a fighter aircraft representative spectrum
(IARPO3a), which equates to approximately one year <strong>of</strong> flying. This is the spectrum applied
to the FT55 fatigue test and FT488/2 bulkhead fatigue test.
<strong>The</strong> crack growth rate and initial discontinuity size were measured for each specimen. This
provided a distribution <strong>of</strong> fatigue initiation starter sizes (“y” intercept in Figure 33) and crack
growth rates (slopes in Figure 33). For a particular stress the fatigue crack growth rate
showed very little variation, therefore only the mean slope was used for all subsequent
analysis. By using the distribution <strong>of</strong> fatigue starters and the mean crack growth rate, Figure
33, it is possible to determine the status <strong>of</strong> each aircraft and the repair procedure needed.
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DSTO-RR-0208
420MPa 390MPa 330MPa
Depth <strong>of</strong> Material to be Removed
(mm)
10.00
1.00
0.10
0.01
0 2000 4000 6000 8000 10000
Simulated Flight Hours (SFH)
Figure 33: A typical plot showing how much material needs to be removed from an un-peened region
to ensure a 1/1000 chance <strong>of</strong> a crack being present. <strong>The</strong> curve is a combination <strong>of</strong> 3σ
fatigue starter and mean crack growth slopes for each stress.
A similar plot for the peened specimens has a higher 3σ initial flaw size but slower fatigue
crack growth rates (ie lower slopes).
<strong>The</strong>se plots form the basis for the repair procedure as they indicate how much material must
be removed from the surface, at a given stress, before the peening process should be
performed. To take full account <strong>of</strong> the peening LIF any “damaged surface material” must be
removed. This ensures that no cracks are present in the surface which would reduce the
effect <strong>of</strong> the repair peening. <strong>The</strong> LIF varies with stress, but at the stresses representing this
region (420MPa) a minimum LIF=1.5 can be expected.
Based on the amount <strong>of</strong> material that could be removed (definite limits based on FE analysis)
and the NDI limits an individual aircraft can be lifed as a safe-life aircraft or a safety-byinspection
aircraft. <strong>The</strong> approach is shown in Figure 34.
43

DSTO-RR-0208
Y470 X19 pocket
Peened
Previously
Non-peened
region
Repair as per
documentation.
Maintain on Safe Life
Determine depth <strong>of</strong> material
to be removed to ensure confident
removal <strong>of</strong> cracks, Figure 33
Based on NDI threshold
and FE analysis.
Requiring less than
500microns for Safe Life
Requiring greater than
500microns for Safe Life
Repair (full cut) as
per documentation.
Maintain on Safe Life
Repair (full cut) as
per documentation.
Maintain on Safe Life
Repair (cut to NDI
limit) Maintain on
Safety-by-Inspection
Figure 34: Summary <strong>of</strong> the proposed DSTO Y470 X19 pocket repair process. In the RAAF fleet the
X-19 pocket has not previously been peened. <strong>The</strong> 500µm (0.02 inch) LPI limit is based on
current AAP NDI limitations.
<strong>The</strong> complete details <strong>of</strong> this repair specification are present in a report by (Sharp, Barter and
Clark, 2000). <strong>The</strong> report allows the RAAF to assess the status <strong>of</strong> each aircraft and determine
future repair times or replacement times.
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3.5.3 IFOSTP Fatigue Tests
In addition to the X-19 repair option described above, the peening process developed by
DSTO and summarised in section 3.3.3 has been used on a number <strong>of</strong> IFOSTP structural test
articles, including FT55, FT46 and FT488/2 bulkhead test. In each case a component/region
was repaired by peening to extend its local fatigue life and allow the full-scale fatigue tests to
continue.
Typical examples <strong>of</strong> structural repairs with the DSTO repair process used on IFOSTP are;
1. FT55 Y470 crotch region repair
2. FT46 stub frame region repair
3. FT488/2 6inch radius repair
Each repair had its own complications but in all cases the damaged material was removed
the surface polished and then the surface peened using the specifications set out in section
3.2.3. <strong>The</strong> peening process has now become a primary option for repair on all IFOSTP tests
for localised life extension.
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DSTO-RR-0208
4. Discussion
<strong>Peening</strong> research around the world and at AMRL has shown that a higher level <strong>of</strong> quality
control is needed for peening aluminium alloys than for high strength steels. This arises
because in both materials, fatigue life is controlled by the residual compressive stress, but for
aluminium alloys the surface damage caused by peening is a major competing effect, tending
to reduce fatigue life. In steels, the residual compressive stress controls fatigue life because
<strong>of</strong> the limited surface damage which can be imparted by the glass beads, and the higher
compressive stresses generated At low fatigue stresses, the growth <strong>of</strong> small surface flaws is
a dominant factor in fatigue life, so large life extension can be achieved by appropriate
peening methods, which will reduce the propagation <strong>of</strong> these small flaws.
This effect is highlighted in early AMRL experiments, Figure 11, which show that a crossover
point exists above which peening can reduce fatigue life depending on the type <strong>of</strong>
peening treatment (ie surface finish). Considering this information, AMRL postulated a
correlation between surface finish (ie roughness) and fatigue life. Further experiments on a
reworked surface supported this correlation except where extensive re-peening <strong>of</strong> damaged
surfaces had occurred. In such cases the surface roughness was low but the fatigue life was
short. Visual examination <strong>of</strong> cross-sections revealed that any debris that was present in the
surface after the original peen, had been hammered deeper into the surface along with the
flattening <strong>of</strong> many laps and folds. This result means that surface roughness does not
necessarily correlate with fatigue life for specimens which have peening directly over
previous peening. High-saturation peening will suffer the same limitations in terms <strong>of</strong>
surface finish/ fatigue life correlation when subjected to these high stresses.
AMRL experiments have clearly demonstrated the competing processes involved in
aluminium alloy fatigue life extension due to peening and have been instrumental in
developing an AMRL rework procedure which can be used to extend the fatigue life <strong>of</strong>
critical areas by repeated polishing and re-peening.
<strong>The</strong>re appears to be very little variation in the residual compressive stresses achieved by
different peening processes. At this stage, only x-ray residual stress measurements have
been made, although a number <strong>of</strong> other residual stress methods are currently being
examined. <strong>The</strong>se include “hole drilling”, neutron radiography and a method developed in
Canada by Marchand. <strong>The</strong> residual stress (240±40MPa) appears to be approximately half the
yield stress (480MPa) <strong>of</strong> 7050 aluminium alloy.
At the stresses tested (410MPa) using load spectrum Spec 16 (RAAF-22010 turning points),
peening is always a benefit to fatigue life compared to a polished finish. However, at higher
stresses poor peening can be detrimental to fatigue life. Early AMRL test results show a
mean LIF <strong>of</strong> between 1.6 to 2.1 due to AMRL rework peening, though this mean life increase
did not fully address the scatter in the peened fatigue life results. If the worst scatter case
were to be used then the LIF would be 1.2. This is significantly lower than the LIF <strong>of</strong> 1.39
suggested by MDA for peening.
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More recent research with the IARPO3a spectrum and normalizing other test results (various
other spectrums) has shown, that for a stress around 400MPa, a LIF=1.5 could be used for the
AMRL rework peening process. This LIF=1.5 accounts for the scatter in the process and
provides the appropriate safety factors needed for DEF-STAN 970 ie. 1/1000 likelihood <strong>of</strong>
failure.
It is also clear from the results that, providing any original peening is removed, a component
can be reworked a number <strong>of</strong> times, restoring the original peened fatigue life each time. Any
reduction in thickness that would occur with each removal <strong>of</strong> the peened surface material
would be course need to be evaluated. One possible limit to this procedure, however, might
be provided by the growth <strong>of</strong> sub-surface cracks to failure; current research is examining this
possibility.
<strong>The</strong>re appear to be a number <strong>of</strong> different peening process combinations in the RAAF and CF
fleet. <strong>The</strong> FS488 bulkhead (or other two wing carry through bulkheads) were delivered in
two conditions.
1. OEM flapper wheel peening or steel bead peening – believed to be on early aircraft.
2. OEM glass bead peening - later aircraft.
As yet AMRL has not been able to determine which bulkheads have experienced which
peening process. Early experiments indicate that steel bead peening process provides severe
surface damage. As yet, AMRL and Canada have conducted no experiments on flapper
wheel peening, which appears to be the process used on all the early bulkheads. To add to
this there have been a number <strong>of</strong> localised rework treatments.
1. ECP45 aircraft -RAAF, polish all original OEM peen finish then re-peen to AMRL
process.
-CF, peen directly over OEM peen.
2. Ceramic Bead Peen - this process is for peening localised areas that generally have not
been previously peened.
This variation in procedures means that the RAAF and CF F/A-18 fleets could have a range
<strong>of</strong> LIF’s depending on the peening process. Clearly, before any allowance for a LIF is used, a
through examination <strong>of</strong> how each peening process variable affects fatigue life is needed.
Additional areas <strong>of</strong> concern are the effect <strong>of</strong> peening into and around radii, leaving unpeened
areas and over-peening. <strong>The</strong>re is also research being performed in Canada
examining a robotic peening system to minimise these variables.
All this work has to date assumed that there is no interaction between microstructure and
peening on fatigue life. AMRL has already shown (Sharp et al. 1996), that there is significant
variation in the microstructure <strong>of</strong> the 7050 bulkheads depending on the date and
manufacturer. As yet, however, no evidence has been seen <strong>of</strong> this having a major influence
on peened fatigue life.
47

DSTO-RR-0208
5. Conclusion
1. <strong>The</strong> benefits <strong>of</strong> peening in terms <strong>of</strong> fatigue life extension are dependent on the applied
stress - the higher the applied stress the lower the benefit. For the test stresses described
in this report, the fatigue life improvement factor for peening ranges from 1.2 to 6 when
using the AMRL proposed rework specification. For a peak stress between 400-420MPa,
common for highly stressed regions on F/A-18 bulkheads, a LIF=1.5 is applicable, which
is slightly higher than the 1.39 LIF declared by the OEM for peening.
2. <strong>The</strong> research undertaken has provided a good understanding <strong>of</strong> the competing effects <strong>of</strong>
residual stress and surface damage in controlling fatigue life in peened components.
3. <strong>The</strong> AMRL research program has developed a rework process which <strong>of</strong>fers the potential
<strong>of</strong> repeatedly restoring the fatigue life <strong>of</strong> critical locations on the F/A-18 aircraft.
4. <strong>The</strong> progressive change in peening media (steel shot to glass bead to ceramic bead) has
caused significant improvements in fatigue life. <strong>The</strong>se changes, combined with a total
loss system or better filtering, has reduced the number <strong>of</strong> broken beads (glass and
ceramic) and odd-shaped beads (steel) that impact the surface. Broadly, this has reduced
the number <strong>of</strong> defects embedded into the surface.
5. Surface roughness (related to peening quality) can be correlated to fatigue life only if the
peening is performed on a clean polished surface - the smoother the peened surface finish
the better the fatigue life. Where peening has occurred directly over earlier peening,
surface roughness cannot be correlated with fatigue life.
6. <strong>The</strong> magnitude <strong>of</strong> the residual compressive stress appears to vary little under the different
conditions examined in this report, although in all cases peening saturation <strong>of</strong> the surface
was ensured. In the case <strong>of</strong> aluminium alloys, the plastically deformed region is many
times larger than the bead size used - this is not the case when peening steel, in which the
deformed region is comparable to the bead size.
7. <strong>The</strong> OEM peening specification does not give the optimal peening conditions for
obtaining the maximum life improvement factor for aluminium alloys. <strong>The</strong> main problem
arises because the peening operation is not mechanised, but relies on human operators,
and the environment is far from ideal in terms <strong>of</strong> making detailed quality control
assessments “on the fly”.
8. Further research is underway and planned to establish a more through statistical basis for
implementing peening methods.
48

DSTO-RR-0208
6. Bibliography
Al-Hassani S.T.S., An Engineering Approach to Shot <strong>Peening</strong> Mechanics, Proc. 2 nd Int. Conf.
On Shot <strong>Peening</strong>, ICSP-2 1984, Editor Fuchs A.O., Paramus, N.J., pp275-282
Al-Hassani S.T.S, Mechanical Aspect <strong>of</strong> Residual Stress Development in Shot <strong>Peening</strong>, 1 st
International Conference on Shot <strong>Peening</strong>, 14-17 September, Paris, Editor Niku-Lari A.,
Pergamon Press, pp 583-602, 1981
Almen J.O, Fatigue <strong>of</strong> Metals as Influenced by Design and Internal Stresses, Surface Stressing
<strong>of</strong> Metals, AM. Soc. Metals, 1947, pp 33-84
Anderson I., Polishing/Glass Bead <strong>Peening</strong> <strong>of</strong> FS488 Bulkhead Rework Area, DSTO minute,
June 1990
Athiniotis N., F/A-18 Fleet 488 Bulkhead <strong>Peening</strong> Pr<strong>of</strong>ile Depths and Microstructural
Deformation, AMRL DAFA Report M40/93, April 1993
Barter S.A. and Houston M.I., Examination <strong>of</strong> Methods to Control the Depth <strong>of</strong> Surface
Removal on the F/A-18 488 Bulkhead, DSTO-DDP-0192, 1997
Barter S.A., FS488 Bulkhead Fracture Surface Preliminary Fractographic Examination, Defect
Assessment and Failure Analysis Report M45/90, July 1990
Barter S.A., Polishing/Glass Bead <strong>Peening</strong> <strong>of</strong> FS488 Bulkhead Rework Area, A/C Mat’s Ref
M59/90/SAB, June 1990
Barter S.A., <strong>The</strong> <strong>Effect</strong> <strong>of</strong> Prior Etching on Fatigue Life, Presentation IFOSTP TRM, Canada,
November 1999
Berry W.R., <strong>Effect</strong> <strong>of</strong> Shot <strong>Peening</strong> with Aluminium Alloy Shot on the Fatigue Life <strong>of</strong>
Extruded L65 Aluminium Alloy, S. and T. Memo 1/60, Ministry <strong>of</strong> Aviation, 1960
Boeing Process Specification BAC5730 RevK, Specification for peening metal parts, 1988
Bresnahan K. and Park J.H., Vertical Stabilizer Stub Frame Life Improvement Testing and
Analysis - Glass Bead Shot <strong>Peening</strong>, Northrop Memo 3881-90-100, 20/6/1990
Bresnahan K. and Park J.H., Vertical Stabilizer Stub Frame Life Improvement Testing and
Analysis - Ceramic Bead Shot <strong>Peening</strong>, Northrop Memo 3881-91-026, 20/6/1990
Butz G.A. and Lyst J.O., Improvement in Fatigue Resistance <strong>of</strong> Aluminium Alloys by Surface
Cold-Working, Materials Research and Standards, ALCOA, December 1961
49

DSTO-RR-0208
Clayton J.Q. and Clark G., <strong>The</strong> <strong>Effect</strong> <strong>of</strong> Steel Shot and Glass Bead <strong>Peening</strong> Treatments on the
Fatigue Resistance <strong>of</strong> 7050-T76351 Aluminium Alloy, Proc. Aust. Fract. Group, “Fracture
Mechanics in Engineering Practice”, Melbourne 1988, pp44-51
Clark G. and Clayton J.Q., <strong>Effect</strong>iveness <strong>of</strong> <strong>Peening</strong> Treatments in Improving Fatigue
Resistance <strong>of</strong> 7050 Aluminium Alloy under Constant Amplitude and Spectrum Loading,
Aust. Surface Engng Conf., University <strong>of</strong> Sth Australia, 1991
Engineering Sciences Data Unit, Guide to the <strong>Effect</strong> <strong>of</strong> Shot <strong>Peening</strong> on Fatigue Strength,
92015, ESDU, May 1992
Fair G, Noble B and Waterhouse R.B., <strong>The</strong> Stability <strong>of</strong> Compressive Stresses Induced by Shot
<strong>Peening</strong> under Conditions <strong>of</strong> fatigue and Fretting Fatigue, Advances in Surface Treatments,
Vol 1, Pergamon Press, UK, 1984
Fuchs H.O., Shot <strong>Peening</strong> <strong>Effect</strong>s and Specifications, ASTM STP 196, Am. Soc. Testing Mats.,
1956, pp22
Gallardo J.L., Optimum Shot <strong>Peening</strong> Intensity using Ceramic Shot on 7050-T7451 5.675-inch
thick Aluminium, Northrop Memo N3881-91-021, 26/2/1991
Graham D., Shot <strong>Peening</strong> <strong>of</strong> the FS488 Bulkhead, ARLENGSTL Report F1/49, January 1986
Hammond D.W. and Meguid S.A., Fatigue Fracture and Residual Stress Relaxation in Shot-
Peened Components, Surface Engineering, Elsevier Science Pty Ltd, 1990, pp386-392
Iida K., Dent and Affected Layer produced by Shot <strong>Peening</strong>, ICSP-2 1984, Paramus, N.J.,
pp283-293
Jaensson B., <strong>The</strong> Influence on the Fatigue Strength <strong>of</strong> Al Alloy parts <strong>of</strong> the relationship
between the Surface Residual Stress State and the Load-Induced Stress State, 1 st International
Conference on Shot <strong>Peening</strong>, 14-17 September, Paris, Editor Niku-Lari A., Pergamon Press,
pp 435-444, 1981
Lambase J., Y488 Bulkhead Fatigue Specimen Failure Analysis, Northrop Memo 3881-90-021,
19/2/1990
Luo W., Noble B and Waterhouse R.B., <strong>The</strong> Interaction between Shot <strong>Peening</strong> and Heat
Treatment on the Fatigue and Fretting Fatigue Properties <strong>of</strong> the High Strength Aluminium
Alloy 7075, Proc 2 nd Int. Conf. on Impact Treatment Procedure, Editor Meguid S.A., Elsevier
Applied Science Publications, Cranfield, September 1986
Luo W., Noble B and Waterhouse R.B, <strong>The</strong> <strong>Effect</strong> <strong>of</strong> Shot <strong>Peening</strong> Intensity on the Fatigue
and Fretting Fatigue Behaviour <strong>of</strong> an Aluminium Alloy, Advances in Surface Treatments and
Surface Finishing, Vol 5, Pergamon Press, UK, 1987, pp 145-153
50

DSTO-RR-0208
Meguid S.A., <strong>Effect</strong> <strong>of</strong> Partial Coverage upon the Fatigue Fracture Behaviour <strong>of</strong> Peened
Components, Fat. Fract. Engng Mater. Struct., Vol 14, No. 5, pp 515-530, 1991
McDonnell Douglas Process Specification, <strong>Peening</strong>, PS 14023 Revision H,
MIL-STD-852, Glass Bead <strong>Peening</strong> Procedures, 18 April 1988
MIL-S-13165C, Shot peening <strong>of</strong> metal Parts, 27 November 1991
Mutoh Y., Fair G.H., Noble B. and Waterhouse R.B., <strong>The</strong> <strong>Effect</strong> <strong>of</strong> Residual Stresses Induced
by Shot <strong>Peening</strong> on Fatigue Crack Propagation in Two High-Strength Aluminium Alloys,
Fatigue Fract. Engng Mater. Struct., Vol 10, No. 4, 1987, pp261-272
Nobre J.P., Kornmeier M., Dias A.M. and Scholtes B., Use <strong>of</strong> the Hole-drilling Method for
Measuring Residual Stresses in Highly Shot-peened Surfaces, Experimental Mechanics, Vol
40, No. 3, 2000, pg 289
Olsoon-Jacques C., Residual Stress Pr<strong>of</strong>ile on Shot Peened F/A-18 aluminium Alloy
Surfaces” DSTO-TN-0230, 1999
Oshida Y. and Daly J., Fatigue Damage Evaluation <strong>of</strong> Shot Peened High Strength Aluminium
Alloy, Surface Engineering, Elsevier Science Pty Ltd, 1990, pp404-416
Roth M and Yanishevsky M, Investigation <strong>of</strong> the <strong>Effect</strong>s <strong>of</strong> Heat Cycles on the Residual
Stresses and Fatigue Life <strong>of</strong> Shot Peened 7050-T7451 Al Alloy, Presentation 21 st IFOSTP
Technical Review meeting, November 1999
Schutz W., Fatigue Life Improvement <strong>of</strong> High Strength Materials by Shot <strong>Peening</strong>, 1 st
International Conference on Shot <strong>Peening</strong>, 14-17 September, Paris, Editor Niku-Lari A.,
Pergamon Press, pp 423-433, 1981
Simpson R.S., Development <strong>of</strong>f a Mathematical Model for Predicting the Percentage Fatigue
Life Increase Resulting from Shot Peened Components, Phase 1, AFWAL-TR-84-3116, April
1985
Sharp P.K. and Clayton J.Q., <strong>The</strong> <strong>Effect</strong> <strong>of</strong> <strong>Peening</strong> on the Fatigue Life <strong>of</strong> Aluminium Alloys,
Technical Presentation DSTO, Melbourne, 1993
Sharp P.K., Clayton J.Q. and Clark G., <strong>The</strong> Fatigue Resistance <strong>of</strong> Peened 7050-T7451
Aluminium Alloy - Repair and Re-Treatment <strong>of</strong> a Component Surface, Fatigue Fract. Engng
Mater. Struct. Vol 17, No. 3, 1994, pp243-252
Sharp P.K. and Byrnes R., <strong>The</strong> <strong>Effect</strong> <strong>of</strong> Mid-Life <strong>Peening</strong> and Rework on Fatigue Life,
DSTO-DDP-164, July 1996
Sharp P.K. and Byrnes R., F/A-18 470.5 Bulkhead Crotch Repair Coupon Tests, DSTO-DDP-
179, July 1996
51

DSTO-RR-0208
Sharp P.K. and Clark G., <strong>The</strong> <strong>Effect</strong> <strong>of</strong> <strong>Peening</strong> on the Fatigue Life <strong>of</strong> Aluminium Alloys, <strong>The</strong>
9 th International Congress on Fracture, Sydney, 1997
Sharp P.K., Barter S.A. and Clark G., Localised Life Extension Specification for the F/A-18
Y470 X19 Pocket, DSTO-TN-0279, May 2000
Shot <strong>Peening</strong> Applications, Metal Improvements Company Inc, Seventh Edition
Smith T.R, FS 488 Bulkhead Shot Peen Coupon Test Program, MDA F/A-18 Project
Presentation, November 1990
Stanier .M., FS488 Bulkhead Polishing During Incorporation <strong>of</strong> AFC44 and AFC129, TFP 399
pt13 (81), August 1990
Ward E.J SQNLDR, Shot <strong>Peening</strong> <strong>of</strong> the Y488 Bulkhead, STL 2506/3/1-2-2 TECH, ,
ENGMECHSTL Report #72, January 1991
Ward E.J SQNLDR, Y488 Bulkhead Shot peening and IVD Coating, STL 2506/3/1-2-2 TECH,
ENGMECHSTL Report #90, May 1991
Ward E.J SQNLDR, Y488 Bulkhead Shot peening and IVD Coating, STL 2506/3/1-2-2 TECH,
ENGMECHSTL Report #98, July 1991
Was G.S. and Pelloux R.M., <strong>The</strong> <strong>Effect</strong> <strong>of</strong> Shot <strong>Peening</strong> on the Fatigue Behaviour <strong>of</strong> Alloy
7075-T6, Met. Trans. A, 10A, May 1979, pp656-658
7. Acknowledgments
<strong>The</strong> authors would like to thank a number <strong>of</strong> people and organisations for their assistance in
the programs discussed in this report.
1. Christina Jacques-Olsson (MPD) and Monash University Department <strong>of</strong> Physics for the
residual stress analysis.
2. ASTA Military Services (Avalon) and Hunter Aviation and Engineering Services
(Williamtown) for specimen preparation and allowing access to their facilities.
3. Rohan Byrnes, Nick Athiniotis and Simon Barter for their work over the years which has
contributed to this report.
4. Dr John Clayton was involved in all <strong>of</strong> the early phases <strong>of</strong> this work.
52

DSTO-RR-0208
Appendix A:
A.1. Spectrum Details
SPECTRUM
NAME
SOURCE TURNING
POINTS
LARGE COUPON
APPLIED STRESS
SMALL COUPON
APPLIED STRESS
Spec 16 RAAF 22010 396 MPa
ST16 USN 5922 410 MPa 417 MPa
IARP03a CF 450MPa, 420MPa,
390MPa, 360MPa
Note for the small specimens K t = 1.035 and for the large coupons K t = 1.055. All sequences
feature repeats <strong>of</strong> are approximately 320 simulated flight hours (SFH).
A.2. Almen Definition
Calibration <strong>of</strong> the impact energy or peening intensity <strong>of</strong> the shot stream is essential for
controlled shot peening. <strong>The</strong> energy <strong>of</strong> the stream is a function <strong>of</strong> the media size, material,
hardness, velocity and impingement angle. In order to specify, measure and calibrate impact
energy, J.O. Almen developed a method utilising SAE 1070 spring steel specimens he called
ALMEN strips.
53

DSTO-RR-0208
Figure 23: <strong>The</strong> standard Almen strips and process <strong>of</strong> measurement (Shot <strong>Peening</strong> Applications 7 th
Edition).
<strong>The</strong>re are three standard Almen strips currently in use; “A” strip 0.051inch thick, “C” strip
0.094 inch thick and “N” strip 0.031inch thick. <strong>The</strong> approximate relationship between the
three Almen strip types is 3N= A= 0.3C. <strong>The</strong> useable range <strong>of</strong> curvature on an Almen strip
is 0.004 to 0.024 inch.
A.3. Saturation Curve
An Almen arc height is not properly termed intensity unless saturation is achieved. In
order to measure Almen saturation, an intensity curve must be developed. Saturation is
defined as the earliest point on the curve where doubling the exposure time produces no
more than ten percent (10%) increase in arc height.
Arc Height
Less than 10% increase
T 2T Exposure Time
Figure 24: <strong>Peening</strong> saturation curve theory (Shot <strong>Peening</strong> Applications, 7 th Edition).
54

DSTO-RR-0208
It should be noted that the saturation <strong>of</strong> the Almen strip and coverage <strong>of</strong> the part are not the
same, nor will they necessarily occur at the same exposure time on the parts being peened.
Saturation is a function <strong>of</strong> the energy <strong>of</strong> the shot stream and establishes the actual intensity
for the particular machine set-up.
Saturation should not be confused with coverage, which refers to the population <strong>of</strong> dimples
on the surface as verified by coverage inspection techniques. Coverage is usually expressed
as a percent coverage (eg 100% coverage). 100% coverage is reached when the original
surface <strong>of</strong> the material is obliterated entirely by overlapping peening dimples. 150%
coverage would be exposing the part to 1.5 times the time required to achieve 100%
coverage. In the case <strong>of</strong> F/A-18, 200% coverage is specified for all peening work.
55

DSTO-RR-0208
A.4. Surface Roughness
<strong>The</strong> measurements <strong>of</strong> surface roughness used in this report are standard measurements
produced by a laser surface pr<strong>of</strong>iler. <strong>The</strong> values are based on German DIN standards but
typical <strong>of</strong> roughness measurements specified in numerous machining handbooks.
Z 1
Z 2 = R max
Z 3
Z 4
L
R max = the largest single roughness depth within the evaluation length. <strong>The</strong> required
standard is DIN 4768
R z = mean value <strong>of</strong> the single roughness depths “Z” <strong>of</strong> consecutive sampling lengths. <strong>The</strong>
required standard is DIN 4768
R z = (Z 1 + Z 2 + Z 3 + ….. + Z n )/n
R a = is the arithmetic mean <strong>of</strong> the areas <strong>of</strong> all pr<strong>of</strong>ile values <strong>of</strong> the roughness pr<strong>of</strong>ile.
R a = [Integral {y(x)}dx] /L
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<strong>The</strong> <strong>Effect</strong> <strong>of</strong> <strong>Peening</strong> on the Fatigue Life <strong>of</strong> 7050 Aluminium Alloy
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18. DEFTEST DESCRIPTORS
Crack propagation, Aluminium alloy, Fatigue, Crack growth, Finishing, <strong>Peening</strong>, Metal finishing
19. ABSTRACT
Many changes in the design and manufacture <strong>of</strong> high-performance military aircraft ⎯ for example, the use <strong>of</strong> highly
optimised design and the use <strong>of</strong> higher-strength material ⎯ have led to an increased sensitivity <strong>of</strong> airframe fatigue life to
surface features such as corrosion or mechanical damage. <strong>The</strong> peening applied to the F/A-18 represents a significant
departure from traditional manufacture, and it is therefore important that the RAAF and AMRL have a through
understanding <strong>of</strong> the peening process, the surface conditions produced, and their effect on structural integrity. This
report discusses the fatigue crack growth research at AMRL, and elsewhere, relating to peening <strong>of</strong> aluminium alloys, and
summarises the improvements in peening which have arisen from this research. . <strong>The</strong> overall aim <strong>of</strong> the peening research
and development discussed was to establish a Life-Improvement-Factor (LIF) for the peening used on the F/A-18, as well
as any future peening required by modifications. It also attempted to provide a means <strong>of</strong> measuring peening quality, to
allow the full exploitation <strong>of</strong> peening to improve fatigue life. It also highlights areas where further research could be
beneficial in relation to peening and the structural integrity <strong>of</strong> the F/A-18 aircraft. <strong>The</strong> report highlights the practical
problems <strong>of</strong> introducing changes to fatigue critical surfaces, with particular reference to the RAAF and CF fleets.
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