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![CHAPTER ONE
Rationale and Organisation
Rhys Jones*, Neil Matthews†
, Alan A. Baker‡,§
, Victor Champagne Jr.¶
*Monash University, Clayton, VIC, Australia
†
RUAG Australia, Bayswater, VIC, Australia
‡
Advanced Composite Structures Australia, Melbourne, VIC, Australia
§
Defence Science and Technology Group, Fairbairn, CBR, Australia
¶
US Army Research Laboratory, Aberdeen, MD, United States
In general, the design of aerospace vehicles requires that all structures and
repairs to load-bearing structural elements be designed in accordance with
damage tolerance design principles, which are largely based on the discipline
of fracture mechanics and which for military aircraft are detailed in the Joint
Services Structural Guidelines-2006 (JSSG2006) [1] and in the USAF Dam-
age Tolerant Design Handbook [2]. This design philosophy has evolved as a
result of a number of high-profile (military aircraft) incidents some of
which,1
together with a number of civilian aircraft failures, are reviewed
in Ref. [3]. As explained in Ref. [3]:
While there have been lessons learned from nearly all of the accidents and inci-
dents, those that were seminal with respect to the design and certification of air-
craft structures were the 1954 Comet failures, the 1958 B-47 accidents, the
1969 F-111 accident, the 1976 AVRO 748 accident, the 1978 Dan Air 707 Lusaka
accident and the 1988 Aloha Boeing 737 accident. The Comet failures resulted
in the use of fail-safe design in commercial aircraft; the B-47 accidents resulted
in the development of the Air Force Aircraft Structural Integrity Programme (ASIP);
the F-111 accident accelerated the introduction of damage tolerance design
requirements in military aircraft; the AVRO 748 and Dan Air Boeing 707 accidents
resulted in complementing commercial fail-safe design with damage tolerance
requirements; and the Aloha Boeing 737 accident brought attention to the impor-
tance that WFD2
has in limiting the safe operation of airplane structures.
The USAF [5] was one of the first to realise that the operational life of an
aircraft was governed by the growth of small cracks and that the analysis tools
1
Ref. [3] also provides a historical perspective as to how the military, the Federal Aviation
Administration (FAA) and industry have dealt with these threats.
2
Widespread fatigue damage (WFD) is defined in Ref. [4] as the simultaneous presence of
cracks at multiple structural locations that are of sufficient size and density such that the
structure will no longer meet the residual strength requirements of FAR section
§25.571(b).
Aircraft Sustainment and Repair © 2018 Elsevier Ltd.
https://doi.org/10.1016/B978-0-08-100540-8.00001-7 All rights reserved.
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![needed for aircraft sustainment issues, that is, to assess in-service cracking,
differed from those used in design. This conclusion was also one of the pri-
mary conclusions reached in the USAF-McDonnell Douglas (now Boeing)
study into sustainment issues associated with cracking in the F-15 aircraft [6]
and is also enunciated in Refs. [7,8]. It also led to the USAF developing a
probabilistic approach to assessing the fatigue life of operational aircraft
[5] in which crack growth was represented as an exponential function, viz.:
a ¼ aieωN
(1)
Here a is the crack length/depth; ai is the initial crack size, which is often
referred to as the equivalent initial flaw size (EIFS) or the equivalent
precrack size (EPS) [9]; N is the number of cycles, or flight hours; and
ω is a constant.
While the operators of military aircraft had long been aware of the
problem of aging of an aircraft [3], public awareness of the problem first
arose via the Aloha Airlines accident in April 1989 [10] which involved
failure due to the presence of multisite damage in the fuselage lap joint
of a Boeing 737-200 aircraft. Indeed, despite the fact that there were
few casualties involved, it continues to be described as the event which first
brought the aging aircraft issue to the attention of the general public and to
those who have responsibility to ensure aircraft safety. The extent of the
problems associated with fuselage lap joints in civil transport aircraft was
further illustrated by the April 2011 incident whereby cracking in a fuselage
lap joint in a Southwest Airlines Boeing 737-300 aircraft resulted in a large
1.52 m (5 ft) hole in the roof [11]. This incident led to the grounding of 79
of Southwest’s older Boeing 737 aircraft [11] and to the cancellation of
almost 700 flights. Subsequent inspections, which found cracks in a total
of four Southwest aircraft [12], led to the US Federal Aviation Adminis-
tration (FAA) mandating the inspection of 175 Boeing 737 aircraft that
had seen more than 35,000 pressurisation cycles. The problem of cracking
in fuselage lap joints was not confined to Boeing 737 aircraft. On 26 Octo-
ber 2010, an American Airlines 757-200 aircraft was forced to land at
Miami International Airport due to a sudden decompression arising from
cracking in a fuselage joint [13]. This aircraft had experienced less than
23,000 cycles. This led to the discovery of cracking in other 757 aircraft
and a subsequent January 2011 FAA Airworthiness Directive [13] mandat-
ing the inspection of all 757-200 and 757-300 aircraft.
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![As a result of these incidents, and other considerations, the FAA intro-
duced the concept of a limit of viability (LOV), defined as the onset of
multisite and/or multielement damage, which the FAA now uses to define
(or limit) the operational life of civil transport aircraft [13,14].
Thus given the central role that fatigue, fracture mechanics and damage
tolerance play in aircraft design and sustainment, the first four chapters of this
book focus on the following:
• The evolution of fatigue design requirements for aircraft structures with
an emphasis on requirements, with design and operational issues also
being addressed (Chapter Two).
• Typical fatigue initiating discontinuities in metallic aircraft structures and
the lead crack concept. (Chapter Three).
• Practical computational fracture mechanics for use in assessing aircraft
structural integrity (Chapter Four).
• Methods for computing/predicting the growth of cracks that arise from
naturally occurring material discontinuities in operational aircraft
(Chapter Five).
Chapters Two–Five are closely linked in that they each discuss the fact that
the growth of ‘lead’ cracks [15], which are defined as the fastest growing
cracks in the airframe and which (as a result) set both the inspection intervals
and the operational life of the fleet, is generally exponential and as such fol-
lows Eq. (1). Chapter Two shows that, in the absence of severe corrosion,
for aluminium alloy airframes the mean value of the EPS (EIFS) is
0.01 mm.
Chapter Five shows that the exponent ω in the exponential growth law is
approximately proportional to the cube of the stress [16]. Chapter Five
explains that this cubic approximation is now used by the Royal Australian
Air Force (RAAF) for the F/A-18 Classic Hornet, the AP3C (Orion) and
the Pilatus Porter PC9 fleets [17–19]. It is further shown that crack growth in
operational aircraft and in full-scale fatigue tests tested under an operational
flight load spectra can be collapsed onto a single ‘master curve’ regardless of
the material or the flight load spectrum.
Chapter Five also shows that, as shown in Ref. [8], the growth of
both long and short cracks can be captured by the Hartman–Schijve variant
of the NASGRO crack growth equation. In this context, it is shown that for
cracks that arise and grow from small material discontinuities in operational
aircraft the Hartman–Schijve equation can be approximated by a simple
Paris crack growth equation with the exponent and the constant of propor-
tionality determined from tests on ASTM E647-13a long crack specimens.
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![The ability of the Hartman–Schijve crack growth equation to compute
accurately the growth of cracks in 7075-T6 P3C Orion test specimen with
intergranular cracking at a dome nut hole that has been repaired using super-
sonic particle deposition (SPD) is further highlighted in Chapter Sixteen.
These findings are important since Chapters Twelve and Thirteen subse-
quently show that a variant of the Hartman–Schijve equation with ΔK rep-
laced by Δ√G, where G is the energy release rate, also captures the growth
of disbonds in adhesively bonded joints and delamination growth in com-
posites [20–23]. As such the Hartman–Schijve equation can be thought of as
a unified formulation that can be used when assessing the original airframe
and the damage tolerance of any subsequent structural repair.
In this context, it should be noted that before the Hartman–Schijve equa-
tion, or any other crack growth equation, can be used to predict/compute
crack growth in an operational structure it is first necessary to determine
the relationship between the stress intensity factor K and the crack length/
depth. To this end, Chapter Four presents simple computational tools, both
three-dimensional (3D) weight function and 3D finite element-based
approaches, for determining the stress intensity factors associated with small
sub-mm cracks in realistic complex geometries under arbitrary flight loads.
Since, as explained in the fatigue test standard ASTM E647-13a, the majority
of the fatigue life of an airframe is generally consumed in growing from a small
sub-mmmaterialdiscontinuitytoasizethatisreadilydetectableusingmodern
nondestructive inspection techniques attention is focused on small initial
cracks that first grow as a complex 3D shape and then transition into a
through-the-thickness flaws with an oblique (part) elliptical crack front.
At this stage, it should be noted that Kinzie and Peeler [24] have shown
how the problem of aging aircraft is particularly acute in the military sphere
and how it has been compounded by US Congress, Public Law 107-314
Sec: 1067 which is entitled: “Prevention and mitigation of corrosion of
military equipment and infrastructure.” One impact of this law is that the
effect of corrosion on the structural integrity of USAF, US Army, US Navy
aircraft and other Navy weapons platforms must be now quantified and
managed as per cracking, that is, in a damage-tolerant manner. In this con-
text, the June 2007 Report to Congress by the Under Secretary of the
Department of Defense (Acquisition, Technology and Logistics) [25] esti-
mated the cost of corrosion associated with US Department of Defense
(DoD) systems to be between $10 billion and $20 billion per year. This
report also outlined the need for research into four primary areas. One of
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![these research areas is: Repair processes that restore materials to an accept-
able level of structural integrity and functionality. The subsequent chapters
address methods that can be used to meet this challenge, viz.: externally
bonded composite repairs, SPD, also known as cold spray (CS) [26], laser
additive deposition (LAD) and multiplicative manufacturing, which
involves the use of both subtractive and additive manufacturing.
In this context, Chapters Six–Thirteen focus on composite repair tech-
nology. This technology has been shown [27–33] to be highly successful in
extending the fatigue lives of both primary and secondary (thin skin) struc-
tures and in ensuring the continued safe operation of aging aircraft.
Chapter Six highlights the key issues associated with bonded composite
repairs and suggests approaches to repairs and reinforcement that contribute
to extending the lifespan and/or the inspection interval when applied to pri-
mary airframe structures. The focus is on the development of alternative
approaches to conventional NDI to assess the structural integrity of the
adhesive bond and based on these approaches the through-life management
of composite repairs.
Chapter Seven focuses on surface treatment and repair bonding and, as
such, complements the work presented in Chapter Six. Both fundamental
and practical aspects of bonded repair application are described.
Chapter Seven also stresses the need for the reproducible durable bonds
which is a fundamental requirement for bonded repair technology.
Chapter Eight details the structural mechanics underpinning composite
repairs to cracks in operational aircraft. Particular focus is given to: the role of
adhesive and matrix nonlinearities; the importance of first ply failure in
design and sustainment; application of the cubic rule in designing composite
repairs and on assessing their effect on operational aircraft; and to the nature
of the test specimens needed for certification.
Chapter Eight also notes that, as explained in the seminal paper by Hart-
Smith [34], adhesive performance is uniquely characterised by the strain
energy density in the adhesive. As such, it clarifies one of the primary design
methodologies enunciated in the PABST report [35], viz. that a critical
design requirement is that the adhesive strain energy density should be kept
beneath its fatigue threshold Wf. A simple means for obtaining Wf, which is
based on the ASTM D1002 thick adherend test, is also presented.
Chapter Eight also shows that for cracks that have grown from small nat-
urally occurring material discontinuities, the role of the fibres bridging the
crack is often a second-order effect. As a result, it is shown that the effect of a
composite patch on crack growth in operational aircraft can be estimated
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![using the exponential crack growth approach presented in Chapter Five, and
in the USAF approach to determine the risk of failure [5], together with the
cubic rule, as also enunciated in Chapter Five, and the assumption that the
effect of the patch is merely to lower the stress field in the underlying struc-
ture. This finding is illustrated both via laboratory coupon tests and by com-
parison with the test programme used to substantiate the boron epoxy repair
of the USAF C-141 aircraft [36] and the repair to the F-111 lower wing
skin [37].
Chapter Nine illustrates a building block approach, which uses the
design methodology outlined in Chapter Eight, to outline the development
of a composite repair to multisite damage in fuselage lap joints similar to
those used in Boeing 737 and Airbus A320 civil transport aircraft. In this
approach, the damage tolerance of the repair is first validated in the labora-
tory and then confirmed by a series of flight demonstrator programmes. For
the repair to the Airbus A320 fuselage lap joint, the damage tolerance of the
repair was validated by application to a full-scale Airbus A320 fatigue test
article. In one demonstrator programme, which was performed as part of
the FAA Aging Aircraft programme [38], nine doublers were applied to a
QANTAS B747-300 aircraft. These doublers experienced a variety of envi-
ronmental conditions ranging from the tropics to Europe in mid-winter.
These doublers were flown for 9 years and saw in excess of 37,000 h
of service and 7020 landings. There was no disbonding or delamination asso-
ciated with these doublers.
Chapter Ten discusses the development of a bonded composite rein-
forcement to the integrally stiffened D6ac steel wing pivot fitting (WPF)
of F-111 aircraft in service with the RAAF. The high load transfer require-
ments made this a very challenging application, especially as it was developed
at an early stage in the development of this repair technology. Particular
attention is again given to a building block approach and as such this chapter
describes the finite element and the experimental programme used to deter-
mine improved stiffener runout geometry as well as the full-scale wing tests
used to validate the computed stress reduction. This repair represents the first
bonded reinforcement to primary structure and was successful in that by
reducing the strains at the critical location by 30%, it achieved its primary
objective namely to stop cracking and wing failure in the RAAF F-111 fleet
and cold proof load test (CPLT). Not only was the WPF reinforcement the
first example of the ability of composite repair technology to eliminate
cracking in primary structure, it was also the first example to highlight
the benefit to be gained by combining composite repair technology with
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![shape optimisation (reworking the local geometry). This approach subse-
quently led to the development of a multiplicative manufacturing solution
to cracking in the LAU7 missile launcher housing (see Chapter Eighteen).
A significant outcome of the F-111 WPF reinforcement programme was
that it resulted in a detailed understanding of the design/assessment tools
needed for designing damage tolerant repairs to meet the stringent damage
tolerant requirements inherent in both JSSG2006 [1] and FAA ac-120-107b
[39]. It is also shown that the in-service performance of the WPF reinforce-
ment reveals the shortcomings in the approach outlined in the US Compos-
ite Materials Handbook CMH-17-3G [40] for the design of bonded joints.
Chapter Eleven presents work performed by the FAA’s Airworthiness
Assurance Center at Sandia National Labs in conjunction with Boeing
and Federal Express, to validate and introduce composite doubler repair
technology to the US commercial aircraft industry. This chapter focuses
on repairs to Boeing DC-10 and MD-11 aircraft and builds on the founda-
tion of a prior successful Lockheed L-1011 door corner repair [41]. The
DC-10 and MD-11 repair programmes highlight the design and analysis,
which is based on the design methodology outlined in Chapters Eight
and Nine, performance (durability, flaw containment, reliability), installa-
tion, and nondestructive inspection issues. This multiyear pilot programme
is particularly important in that it was perhaps the first to demonstrate the
successful and routine use of composite doubler repairs on commercial
aircraft.
In this context Chapter Eleven, when taken together with the work
presented in Chapters Eight–Ten, further substantiates the extended PABST
design methodology [35] whereby the adhesive strain energy density should
be kept beneath its fatigue threshold Wf.
In 2009, the US FAA introduced a slow growth approach for certifying
composite and adhesively bonded structures [39]. This approach requires
that delamination and/or disbond growth under cyclic-fatigue loads be
slow, stable and predictable. To this end, Chapter Twelve addresses the chal-
lenge of developing a methodology capable of enabling this slow growth
approach to certification to be implemented. As a result Chapter Twelve
examines the growth of disbonds from small naturally occurring material dis-
continuities in adhesively bonded structures. It is shown that, for the exam-
ples studied, the disbond growth histories can be accurately computed using
a form of the Hartman and Schijve variant of the NASGRO crack-growth
equation discussed in Chapter Four with ΔK replaced by Δ√G. It is also
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![shown that the scatter in the disbond growth histories can be captured by
allowing for small changes in the fracture-mechanics threshold term
Δ√Gthr. These findings suggest that the Hartman–Schijve crack-growth
equation has the potential to address the ‘slow growth’ approach to certify-
ing composite/bonded structures and bonded repairs outlined in the US
FAA Airworthiness Advisory Circular No: 20-107B [39].
The growth of disbonds in adhesively bonded repairs represents only one
of the possible failure mechanisms. Externally bonded composite repairs can
also fail by delamination. Indeed, the growth of delaminations in polymer–
matrix fibre composites under cyclic-fatigue loading in operational aircraft
structures has always been a very important factor. It has the potential to
affect significantly the service life and is central to the damage tolerant assess-
ment of both composite repairs and composite airframes [1,35,42]. Until
recently, the recommended approach was to adopt a ‘no growth’ design
[40]. However, the introduction by the FAA of a ‘slow growth’ approach
to the certification of composite and bonded structures [39] has focused
attention on the experimental data, the analytical tools needed to assess
the growth of delaminations under fatigue loads and on means for determin-
ing a lower bound on the fatigue threshold. As such the focus of
Chapter Thirteen is to address the topic of the growth of delaminations
in polymer–matrix fibre composites under cyclic-fatigue loading using a
fracture mechanics approach. Specific attention is given to the test and
data-reduction procedures that are required in order to determine a ‘valid’
rate of crack growth, da/dN, versus the range of the energy release rate, ΔG
(or the maximum strain-energy release rate, Gmax, in a cycle) relationship.
This is particularly important since only such ‘valid’ relationships can be reli-
ably used (a) to characterise and compare different types of composites and
adhesives and (b) for lifting the in-service behaviour of composite structures
and bonded composite repairs. To this end, Chapter Thirteen highlights the
important role played by fibre bridging that may occur behind the tip of the
advancing delamination, and which causes retardation of the fatigue crack
growth rate. It is shown that such retardation effects cannot usually be
avoided when using the Mode I double-cantilever beam test to ascertain
experimentally the fatigue behaviour of composites. Consequently, a means
of estimating a ‘valid’ (low retardation) relationship is needed.
Chapter Thirteen presents one such approach that is based on the use of
the Hartman–Schijve delamination growth equation discussed in Chapters
Five and Twelve. This approach also enables a lower bound on the fatigue
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![threshold to be determined. This ‘lower bound’ is particularly important if
the design is to be based on a no growth approach to delamination or
disbonding.
Additive manufacture (AM) is increasingly being referred to as the next
industrial revolution in that 3D complex shapes can be built layer by layer
through material deposition. AM is fast approaching a ‘mainstream’ technol-
ogy status [43]. In the aerospace industries, companies that are expanding
their 3D manufacturing and production engagement are general electric
for jet engines and Lockheed Martin and Boeing for aerospace and defense
[43]. Additive metal technologies (AMTs) are part of the additive
manufacturing family. In aerospace, AMT can be used for a range of
sustainment applications as well as for the creation of replacement parts.
In this context, Chapters Fourteen–Eighteen focuses on: SPD, also known
as CS, LAD and multiplicative manufacturing.
To help clarify the basics of AMT, Chapter Fourteen introduces SPD.
SPD, also known as ‘cold spray (CS)’, is a materials consolidation process
whereby micron-sized particles of a metal, ceramic and/or polymer are
accelerated through a spray gun fitted with a De Laval rocket nozzle using
a heated high-pressure gas (i.e., helium or nitrogen) such that the particles
exit at supersonic velocities and consolidate on impacting a suitable surface
to form a coating or a near-net-shaped part by means of ballistic impinge-
ment [44–49]. The particles utilized are typically in the form of commer-
cially available powders, ranging in diameter from about 5 to 100 μm,
and are accelerated at velocities from 300 to 1500 m/s by injection into a
high-velocity stream of gas. Particles that are 5 μm in diameter do not have
enough momentum to leave the gas stream and impact the substrate. Hence,
nanoparticles cannot be deposited using CS in this manner. Instead agglom-
erated nanoparticles or nanostructured powder within the feasible size range
are used and CS has been shown to produce nanostructured coatings, as well
as nanostructured bulk materials [50]. Chapter Fourteen also briefly discusses
the advantages and potential of SPD for both geometry restoration and as a
structural repair.
Chapter Fifteen builds on this introduction to SPD. It notes that for the
new and next generation aircraft, utilisation of computer-aided design tools
and technologies and the concurrent computer-aided manufacturing pro-
gramming of highly sophisticated machines will often result in optimised
designs that significantly reduce and even eliminate any excess material to
reduce the overall aircraft weight and thereby enhance operational perfor-
mance. Chapter Fifteen further remarks that one outcome of this ‘optimised’
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![design process is that these new and next generation components will often
not be able to be repaired/recovered through the conventional subtractive
processes (blending) nor will they be able to be recovered using external
mechanically fastened repairs. Hence the challenge is to develop additive
metal solutions so that components and/or structure can be returned to
operational service in a safe, timely and cost-effective manner.
In this context, Chapter Fifteen notes that the maintenance of an aircraft
fleet is an ongoing and evolving process and involves both the repair of leg-
acy aircraft and the development of capability for next generation aircraft.
For legacy aircraft maintenance procedures identify components and struc-
ture that no longer meet specification or are considered unfit for purpose and
are typically replaced or require extensive repairs. In many instances, the fail-
ure to meet specification is due to the loss of metal from the component or
structure. This loss can occur through corrosion, wear, erosion or impact.
These replaced components or repaired structures can have a significant
price tag and adversely affect aircraft availability and as such contribute sig-
nificantly to the cost of ownership of the fleet.
In the military sphere it has already been noted that the June 2007 Report
to Congress by the Under Secretary of the DoD (Acquisition, Technology
and Logistics) [25] estimated the cost of corrosion associated with US DoD
systems to be between $10 billion and $20 billion annually. This report out-
lined the need for research into four primary areas one of which was: Repair
processes that restore materials to an acceptable level of structural integrity
and functionality. In a recent independent report, the annual cost of corro-
sion (2013) for the Australian Defense Aerospace sector was quoted as
A$228 million [50].
To meet this challenge Chapter Fifteen focuses on two AMTs, namely
SPD and LAD, which are currently being used in the sustainment of aero-
space vehicles to reduce the cost of ownership and increase operational
availability.
Having introduced the concept of SPD (CS) and how it can be can be
used to restore geometry and having noted that in the past decades, partic-
ularly since the Aloha Airlines incident 1989 [3], Chapter Sixteen notes that
corrosion has become one of the primary considerations in both military and
commercial aircraft. Indeed, numerous authors, for example, Refs. [51–54],
have noted that corrosion can impact heavily the economics, maintenance
and safety of aircraft fleets.
A number of different types of corrosion have been detected on aircraft
structural aluminium alloys with exfoliation, pitting, intergranular corrosion
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![(IGC) and stress vorrosion cracking (SCC) being the prominent types of
corrosion generally occurring in structural components exposed to corrosive
environments [52,54–56]. This corrosive degradation can result in a signif-
icant adverse impact on structural integrity through the loss of material and/
or the generation of cracking. Consequently, before discussing the ability of
SPD to restore structural integrity Chapter Sixteen presents a brief discussion
on corrosion in aircraft structures.
Chapter Sixteen then shows how AMT, specifically SPD (CS), can be
used to alleviate the effect of environmental degradation on the structural
integrity of wings and horizontal stabilisers. Particular attention is paid to
(a) The ability of SPD to essentially restore the load carrying capacity of
upper wing skins with skin corrosion and rib stiffened upper wing
planks with SCC in the risers.
(b) The ability of SPD repair corrosion on the lower (wing skin) tension
surfaces.
(c) The ability of SPD to alleviate the deleterious effect of IGC at fastener
holes on the lower (wing skin) tension surfaces.
(d) The ability of the Hartman–Schijve crack growth equation, which is
discussed in Chapter Four, to capture the growth of cracks in structures
repaired using SPD.
Chapter Seventeen builds on the work presented in Chapter Sixteen. To this
end, Chapter Seventeen summarises a building block approach that has been
used to demonstrate the potential of SPD for repairing and enhancing the
airworthiness and integrity of fuselage skins, fuselage lap joints and centre
barrel structures. The initial focus is on the application of SPD doublers
to fuselage lap joints representative of a Boeing 737 civil transport aircraft
and as such complements the work presented in Chapter Nine, which dis-
cussed composite repairs to Boeing 737 fuselage lap joints. These successful
tests are then supported by an application to an F/A-18 Hornet wing attach-
ment centre barrel laboratory test article. It is noted that the centre barrel was
subsequently tested for more than three design lifetimes without detectable
evidence of degradation (disbonding, delamination or cracking) to the SPD
doublers, including those over fastener holes, or the surrounding structure.
Chapter Seventeen also shows that even though, during fastener removal to
install a Canadian modification mid-way through the test, the SPD was
damaged there was no cracking, delamination or disbonding induced and
the SPD, and the surrounding structure, remained intact even after cycling
for a further 13,000 flight hours, which corresponds to in excess of 2 design
lifetimes. This finding highlights the damage tolerant nature of SPD. As such
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![Chapter Seventeen shows SPD to be highly durable and damage tolerant and
that it can be used to restore structural integrity of full-scale aircraft compo-
nents under a representative fighter aircraft spectrum. Chapter Seventeen
also shows that SPD doublers can withstand cyclic stress amplitudes of more
than 200 MPa without disbonding, delamination or cracking. This is an
important finding since it represents a realistic upper bound on the stresses
seen on the exterior wetted surface of many operational aircraft.
Whereas Chapters Fourteen–Seventeen focused on AMT we have seen
that Chapter Ten addressed the use of structural optimisation in conjunction
with an externally bonded composite repairto addresssuccessfully theproblem
of cracking in primary structure. The success of this combined approach sug-
geststhattoaddressproblemsassociatedwithaircraftsustainmentitmaybepos-
sible to use multiplicative manufacturing, which is defined as the simultaneous
application of subtractive and additive manufacturing processes, to extend the
fatigue life of an operational aircraft. To this end, Chapter Eighteen illustrates
the potential of multiplicative manufacturing via the LAU-7 missile launcher
housing [57]. In this context, Chapter Eighteen show that applying SPD to
the external surface of the LAU-7 missile launcher housing further reduces
the peak stress at the critical location, in comparison to only adopting a shape
optimizationsolution,by29%.Itisshownthatthisadditionalreductioninthe
stress field would enable the allowable crack size to be increased from 0.5 to
2.14 mm and would effectively triple the time to the allowable crack size.
As such Chapter Eighteen illustrates the potential for using multiplicative
manufacturing to extend the fatigue life and maintain the capability of opera-
tional aircraft.
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[23] A. Mujtaba, S. Stelzer, A.J. Brunner, R. Jones, Influence of cyclic stress intensity thresh-
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ings Navy Corrosion Conference, San Diego, 7th–10th, June, 2010.
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![[28] A.A. Baker, L.R.F. Rose, R. Jones, Advances in the Bonded Composite Repair
of Metallic Aircraft Structure, Vols. I and II, Elsevier, Cambridge, MA, ISBN 0-08-
042699-9, 2002.
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Aircraft Structure, Elsevier Science, Oxford, 2002, pp. 871–884.
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aircraft using BFRP crack patching technology, Theor. Appl. Fract. Mech. 2 (1)
(1984) 1–16.
[32] A.A. Baker, R.J. Chester, M.J. Davis, J.D. Roberts, J.A. Retchford, Reinforcement of
the F-111 wing pivot fitting with a boron/epoxy doubler system—materials engineer-
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F-111C wing pivot fitting: structural aspects, Compos. Struct. 11 (1) (1989) 57–83.
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J. Mazza, Bonded boron patch repair evaluation: final report, USAFA-TR-2007-06,
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[37] K.F. Walker, L.R.F. Rose, Case history F-111 lower wing skin repair substantiation
(Chapter 27), in: A. Baker, L.R.F. Rose, R. Jones (Eds.), Advances in the Bonded
Composite Repair of Metallic Aircraft Structure, Elsevier, Cambridge, MA, ISBN
0-08-042699-9, 2002.
[38] Atluri SN., Park JH., Punch EF., O’Donohue PE. and Jones R., Composite repairs of
cracked metallic aircraft, Federal Aviation Administration, contract report, DOT/
FAA/CT-92/32, May, 1993.
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aircraft (Chapter 28), in: A. Baker, L.R.F. Rose, R. Jones (Eds.), Advances in the
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2002. 0-08-042699-9.
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management strategy, ADF Aircraft Structural Integrity Symposium (AASIS16), 16th–
18th November, 2016. Available from: http://www.defence.gov.au/DASP/
ConferencesAndSeminars/Presentations.asp.
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ings, J. Therm. Spray Technol. 11 (4) (2002) 542.
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blog/cold-spray-ge-technology. GE Global Research, 2013. 07 November.
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Assisted Cracking in Aging Military Vehicles, 2011, pp. 27-1–27-42, RTO-AG-AVT-
140, Research and Technology Organisation of NATO, March.
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FURTHER READING
[1] https://data.epo.org/publication-server/rest/v1.0/publication-dates/20131002/
patents/EP2644510NWA1/document.pdf.
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![QZ ground loads
RAAF Royal Australian Air Force
R.H. right hand
RUL remaining useful life
SHM structural health monitoring
S–N stress versus number of cycles
USAF United States Air Force
VA variable amplitude
WFD widespread fatigue damage
XWB eXtra Wide Body
ΔP pressurisation differential
ε–N strain versus number of cycles
1 EVOLUTION OF FATIGUE REQUIREMENTS
Since the early 1930s the evolution of aircraft structural integrity has
been concerned largely with the service behaviour of high-strength metallic
materials, particularly aluminium alloys. Taking a (very) broad view, the his-
tory of this evolution is as follows [1,2]:
1930–40: Development of all-metal aircraft for public transport. Design
and analysis emphasised static strength, with little consideration of fatigue.
1940–56: Increasing awareness, notably for civil aircraft, of the impor-
tance of fatigue for airframe safety. Materials with higher static strengths
were developed. However, aluminium alloys did not - and still do not -
show corresponding increases in fatigue strength. Design of civil aircraft
became based on both static and fatigue strengths (safe-life fatigue design)
in 1956.
1956-today: Development of Fail-Safe and Damage Tolerant design
methods for both civil and military aircraft. These methods recognise that
airframe structures must withstand service loads even when damaged or
cracked. Safety should be ensured by fatigue and fracture testing and analysis
of damaged structures; preservice and in-service inspections; and eventual
repairs, replacements or retirement.
Service failures have greatly influenced these developments. In particu-
lar, there are six case histories that may be regarded as ‘milestones’ in the
aerospace industry’s approach to structural integrity [3–7]. By happenstance
these case histories are evenly divided between civil and military aircraft, see
Figs 1 [2,8] and 2 [2,9,10]. The significances of these accidents are discussed
in Sections 1.1 and 1.2.
18 Russell J.H. Wanhill](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-26-2048.jpg)
![1.1 Civil Aircraft Milestone Accidents
De Havilland Comets: In January and April 1954 two Comets suffered pres-
sure cabin disintegration resulting from fatigue. Investigation and follow-up
gave a general awareness of finite fatigue lives, the significance of Fail-Safety,
and the usefulness of full-scale fatigue testing (FSFT).
Fail-Safe design principles mean that major parts of the structure are
(were) designed firstly to achieve a satisfactory fatigue life with no significant
cracking. Secondly, the structure is (was) also designed to be inspectable in
service and able to sustain significant and easily detectable damage before
1950
COMETS
1954
DAN-AIR
1977
ALOHA
1988
1960 1970
Fatigue (Safe-Life)
Either
Fail-Safe
Inspection impractical?
Yes
No
Damage tolerance
Full-scale fatigue test
LOV*
1980 1990 2000 2010
Fig. 1 Milestone civil aircraft accidents and evolution of civil aircraft fatigue require-
ments [2,8]: *LOV ¼ limit of validity.
1950
B-47’S
1958
F-111
1969
MB326H
1990
1960 1970
Fatigue (Safe-Life)
Static strength
requirements
only
USAF ASIP*
Damage tolerance
USAF
Both
Quantitative fractography (QF) and
ASIP*: DSTO/RAAF
1980 1990 2000 2010
Fig. 2 Milestone military aircraft accidents and the evolution of military aircraft
fatigue requirements [2,9] and techniques [10]: *ASIP ¼ aircraft structural integrity
program(mes).
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![safety is compromised. These latter requirements were met mainly by struc-
tural design concepts having multiple load paths and established residual
strength (fracture) requirements in the event of complete or obvious partial
failure of one structural element.
These Fail-Safe design principles were adopted in 1956, with some
exceptions, notably including landing gears, which are not amenable to
Fail-Safe design. However, as yet there was no requirement to conduct FSFTs,
see Fig. 1. Although FSFT has generally or often been done since the Comet
accidents, the requirement for FSFT became mandatory only in 1998, see
the last paragraph about the ALOHA Boeing 737 accident.
DAN-AIR Boeing 707: In May 1977 a Boeing 707 freighter lost the R.
H. horizontal stabiliser owing to fatigue in a supposedly Fail-Safe spar. There
were several contributing factors:
• inadequate inspection for detecting partial failure of the spar upper chord
• inadequate residual strength after complete failure of the upper chord
• unanticipated high service loads
• stabiliser modification not checked by FSFT.
This accident prompted (i) reconsidering the problems of ageing aircraft: it
became clear that some inspection methods and schedules were inadequate
and required supplementary programmes; and (ii) the need for damage tol-
erance (DT), which replaced the Fail-Safe design principles in 1978.
In civil airworthiness terms, DT means that fatigue cracks will be
detected before the safety is compromised, but unlike Fail-Safety there is
no premise that cracking will become obvious before it reduces the residual
strength below the required safety level [8].
ALOHA Boeing 737: In April 1988, a Boeing 737 suffered explosive
decompression with a loss of part of the pressure cabin, subsequently landing
safely. The physical manifestation of the accident was multiple site fatigue
damage (MSD) along a critical rivet row of the upper skin lap splice. There
were several contributing factors:
• skin splice design relied on cold-bonding as well as riveting
• bonds susceptible to corrosion and disbonding
• neglected inspections and repairs.
This accident led to recognition that widespread fatigue damage (WFD), of
which MSD is one aspect, can cause loss of Fail-Safety. It also prompted
more actions to ensure the safety of ageing aircraft, and corrosion control
programmes for both civil and military aircraft.
Other instances of WFD led eventually to mandatory FSFT from 1998
onwards; and in 2011 the limit of validity (LOV) concept for aircraft above
20 Russell J.H. Wanhill](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-28-2048.jpg)
![34,000 kg. The LOV concept requires FSFT to determine the onset of
WFD [11], and as such effectively sets a Safe-Life equivalent for this category
of aircraft.
1.2 Military Aircraft Milestone Accidents
Boeing B-47s: In March and April 1958 four Boeing B-47s suffered fatigue
failures resulting in wing losses. After the fourth accident a fleet recovery
programme was initiated. This required development of a structural integ-
rity programme that included FSFT. This programme was completed
in 1959.
Also in 1959, the U.S. Air Force (USAF) adopted Safe-Life fatigue
design principles and a permanent aircraft structural integrity program
(ASIP), see Fig. 2, that included FSFT. This became a formal requirement
(mandatory) in 1969.
It should be noted that all ASIPs, including those introduced by the
Royal Australian Air Force (RAAF) after the MB-326H accident in 1990
(see below), involve much more than fatigue issues [12].
General Dynamics F-111: In December 1969 a General Dynamics
F-111A lost the left wing after only 107 airframe flight hours and at less than
half the design limit load. Failure occurred by fast fracture after a small
amount of fatigue from a large manufacturing flaw in the lower plate of
the L.H. wing pivot fitting. The plate was made from a high-strength steel
with limited fracture toughness.
This failure and FSFT fatigue and fracture problems led to a fracture con-
trol programme for the F-111 critical steel components [13]. Furthermore,
the F-111 problems and early fatigue cracking and WFD in Lockheed C-5A
wing boxes [14] resulted in the USAF abandoning the Safe-Life policy and
introducing their DT approach in 1974–75 [15,16]. This preceded the civil
aircraft DT requirements by 3–4 years, cf. Figs 1 and 2. There are some dif-
ferences, the main one being that the USAF DT approach requires that initial
flaws or cracks should be assumed to be present in new structures.
However, there is no universal acceptance of the USAF DT require-
ments for military aircraft. For example, the U.S. Navy still uses the Safe-
Life approach.
Aermacchi MB-326H: In November 1990, an Aermacchi MB-326H lost
the left wing owing to fatigue failure in the wing spar at only 70% of the
Safe-Life derived from FSFT. Investigation showed that fatigue began from
a badly drilled bolt hole.
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![Further investigation revealed (i) manufacturing flaws in many holes,
(ii) fatigue cracks growing from bolt hole manufacturing flaws in spars from
other aircraft, and (iii) fatigue cracks growing from normal quality holes and
other structural details [17].
A fleet recovery programme based on teardown of several high-life
wings, followed by quantitative fractography (QF) of the growth of in-
service fatigue cracks, showed that the original FSFT-based Safe-Life was
highly unconservative and that wing replacements were the only feasible
option.
In addition the RAAF’s structural integrity policy changed similarly to
the USAF’s after the B-47 accidents in 1958. Since 1990 the RAAF has
established comprehensives ASIPs for each fleet. Some have included
QF-based fatigue analyses.
2 CONTINUING DEVELOPMENTS
This section summarizes some developments in fatigue analyses and
airframe materials that have implications for future aircraft structural integ-
rity requirements and ASIPs.
2.1 Fatigue Analyses for Conventional Alloys
Ongoing fatigue research aims to improve structural analysis capabilities and
methods for fatigue life and crack growth predictions. Two of the main
developments are (i) the study of short fatigue crack growth and (ii) the
possible and actual effects of corrosion on fatigue and the combined action
of corrosion and fatigue.
Short crack growth: When the USAF DT approach was first introduced,
in the early 1970s, there were few actual data for short crack growth. Owing
to advances in QF this is no longer the case, particularly since the MB326H
accident in 1990. There are now many data for early crack growth in high-
performance aircraft [10], including the types and sizes of fatigue-initiating
initial discontinuities [18]. Analysis methods have been developed to use
these data for life predictions and reassessments, including the use of equiv-
alent precrack sizes (EPS) to represent realistic initial discontinuities [10].
Corrosion and fatigue: Since the ALOHA Boeing 737 accident there has
been much effort to develop life-prediction models combining corrosion
and fatigue. However, most of these models have been based on laboratory
22 Russell J.H. Wanhill](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-30-2048.jpg)
![testing rather than service experience, and have assumed that corrosion
would not only initiate fatigue but also accelerate it.
Evidence from military aircraft indicates that although corrosion can ini-
tiate fatigue, the corrosion contribution is predominantly ground-based,
while fatigue occurs mainly during flight [19,20].
At present the question whether corrosion and fatigue act in combina-
tion or independently requires further investigation per aircraft type and
operational conditions. This issue is important since it can have a major
impact on service life management [19,21].
2.2 Airframe Materials
Aluminium alloys have predominated in airframes since introduction of the
Boeing 247 (1933) and Douglas DC-2 (1934), but composites (mainly car-
bon fibre reinforced plastic, CFRP) and titanium alloys provide competition
for certain applications. It is also important to note that the choices of mate-
rials for different types of aircraft vary greatly:
Transport aircraft: Aluminium alloys still account for about 60% of the
airframe structural weight, with the notable exceptions of the Boeing 787
Dreamliner and Airbus A350 XWB, which use 50%–53% composites and
only 19%–20% aluminium.
Tactical aircraft: The emphasis tends to be on composites and lesser per-
centages of titanium, aluminium and steels. For example, the British Aero-
space/DASA/CASA/Alenia Eurofighter and Lockheed Martin F-35 both
use 35%–40% composites, most of which are in the external skin and
substructure.
There are exceptions, notably the Lockheed Martin and Boeing F-22,
which uses only 24% composites and 20% aluminium. Much of the F-22
structure (42%) is titanium, which can tolerate the relatively high service
temperatures, especially in the aft fuselage.
Helicopters: Owing to the importance of weight savings for vertical lift
aircraft, even higher percentages of composites are being used: e.g. the Bell-
Boeing V-22 Osprey airframe consists of nearly 80% composites [22] and the
NHIndustries NH90 airframe is over 90% composites.
The modern trend is to design and build hybrid airframe structures, not
least because of the availability of 3rd-generation aluminium–lithium
(Al–Li) alloys. Al–Li alloys provide practical structural efficiencies that
can rival those of CFRP composites, resulting in similar weight savings with
respect to conventional aluminium alloys. The Airbus A380 represents an
important example of a recent hybrid airframe construction, see Fig. 3.
23
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![The structural integrity implications of using the newer materials are
summarised here:
(1) Replacement of conventional aluminium alloys by the newer (3rd)
generation Al–Li alloys will hardly change the fatigue and fracture
structural analysis approaches, since these alloys have reached a similar
level of reliability in their engineering properties [23].
(2) GLARE (glass reinforced aluminium laminates) is a composite used
extensively in the Airbus A380 pressure cabin, see Fig. 3. GLARE
has fatigue and fracture properties amenable to conventional analyses
for all-metal airframes, although there are some differences. These have
led to a GLARE design ‘toolbox’ to enable compliance with current
airworthiness regulations for metallic structures [24].
(3) CFRPs have very different properties and design principles. They
have high fatigue resistance when free from defects and stress con-
centrations, but they are susceptible to impact damage and subse-
quent cracking and delamination. The damage growth—which need
not be by fatigue—is difficult to predict. This also makes it difficult
to validate repairs. Other related problems are inspection reliability
and difficulties in analysing complex components to predict the onset
of failure.
GLARE upper
fuselage panels
Al-Li 2099-T83,
2196-T8511 and
CFRP floor beams
CFRP empennage
CFRP pressure
bulkhead
Newer AA 7XXX (upper wing) and 2XXX
(lower wing) alloys; Al-Li 2050 –T84
internal structure; CFRP ribs
AA 6XXX alloy LBW
lower fuselage panels
CFRP centre
wing box
Fig. 3 Advanced materials and processing for major structural areas of the Airbus
A380 [23]: CFRP ¼ carbon fibre reinforced plastic; GLARE ¼ glass reinforced alumi-
nium laminates; AA 2XXX, 6XXX and 7XXX ¼ conventional aluminium alloys;
Al-Li ¼ aluminium–lithium alloys; LBW ¼ laser beam welding.
24 Russell J.H. Wanhill](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-32-2048.jpg)
![The overall result is that safety is currently ensured by over-
designing CFRPs according to the ‘no growth’ DT principle
[25,26]. However, this does not mean that there are no further devel-
opments. The Federal Aviation Administration (FAA) actively sponsors
research in several key areas, including structural substantiation, DT and
maintenance practices, materials control and standardisation, and
advanced material forms and processes [27]. These activities continually
update the safety and certification requirements for composite struc-
tures [27,28].
Looking to the future, there is a current lack of FSFT programmes
for CFRP components under realistic conditions. These include the
effects of temperature and humidity on the properties.
3 FATIGUE LIFING METHODS FOR METALLIC
AIRFRAME STRUCTURES
The evolution of fatigue requirements discussed in Section 1 and illus-
trated by Figs 1 and 2 has meant that several fatigue lifing methods have
evolved, see Table 1. Some comments on the methods are made here.
Stress-life (S–N): This method has been used extensively in the Safe-Life
approach. The method relies on simple cumulative damage rules to obtain
‘safe’ fatigue lives under constant amplitude (CA) or variable amplitude (VA)
load histories. The safe lives are then factored down using scatter factors
based on engineering judgement and experience.
Strain-life (ε–N): This is a common method used to analyse fatigue dam-
age in aircraft fleets, and there are several Original Equipment Manufacturer
(OEM) and publicly available tools for doing this.
USAF DT: This method is used for many USAF aircraft. It has been
effective in ensuring structural safety, whereby the assumed initial flaw or
crack sizes (see the F-111 discussion in Section 1.2) are large enough for frac-
ture mechanics calculations of FCG using models based on well-established
macrocrack growth behaviour.
This DT approach has been less satisfactory for analyses of structural dura-
bility, i.e. achieving an economic life based on FCG. The initial flaw sizes for
durability analyses (equivalent initial flaw sizes, EIFS) were set at 0.254 mm for
fastener holes (the most critical locations), with smaller sizes allowed by dem-
onstration. These initial crack sizes are well within the short crack regime, where
until recently there were few relevant data and reliance had to be made on
back-extrapolation of long crack growth behaviour using ‘traditional’ analyt-
ical modelling. These issues have been addressed by the DSTO approach.
25
Fatigue Requirements for Aircraft Structures](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-33-2048.jpg)
![Table 1 Survey of Methods for Fatigue Life Assessment of Metallic
Airframe Structures [10]
• Stress–life (S–N)
fatigue limits, Se; unnotched and notched (Kt); constant amplitude (CA) data
modifications to Se
mean stress effects (R)
linear damage rule, also for variable R
scatter factors
• Strain–life (ε–N)
strain-life equation, unnotched data, R ¼ 1
cyclic stress–strain curve analysis
rainflow cycle counting (closed hysteresis loops)
stress–strain at critical location (notch analysis)
mean stress effects (R) via equivalent strain equations, leading to equivalent
strain amplitudes
damage accumulation rule
• USAF Damage Tolerance (DT)
specified equivalent initial flaw sizes (EIFS) based on NDI capabilities
back-extrapolation of long crack growth data to derive short crack growth
LEFM long crack growth models (noninteraction, yield zone, crack opening,
strip yield) to derive variable amplitude (VA) crack growth from constant
amplitude (CA) data
possible use of the crack opening model for short cracks (FASTRAN);
differences in long and short crack thresholds need to be included
mainly deterministic: stochastic approach becoming accepted
• DSTO approach: implemented by the RAAF
actual initial discontinuity/flaw sizes and their EPS
actual short-to-long crack growth data using QF
data compilations to establish empirical relationships describing crack growth
behaviour
deterministic (‘upper bound’) estimates of lead crack growth
scatter factors
• Holistic approach: proposed
fatigue nucleation mechanisms (also as functions of notch stress
concentrations, Kt)
fatigue nucleation lives (S–N and/or ε–N assessments)
evaluation and selection of marker load strategies for QF of short-to-long
crack growth
actual short-to-long crack growth using marker loads and QF
establishment, validation and choice of appropriate crack growth models and
‘laws’
deterministic (‘upper bound’) and stochastic approaches
environmental effects, notably corrosion
26 Russell J.H. Wanhill](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-34-2048.jpg)
![DSTO approach: As mentioned in Section 2.1, advances in QF since
1990 have provided many data for early crack growth in high-performance
aircraft. Analyses have been developed to use these data for FCG life
predictions [10].
Holistic approach: This is a broad programme, see Fig. 4, and is intended
for all aircraft types. The inclusion of corrosion is especially noteworthy: see
also the remarks about corrosion and fatigue in Section 2.1. The project
sponsor is the Canadian National Research Council (NRC), with contribu-
tions from other experts from North America, Europe, Australia and Asia
( Japan).
4 FATIGUE LIFING ANALYSES FOR METALLIC
AIRFRAME STRUCTURES
This section is intended to give generic overviews of fatigue lifing ana-
lyses. Much more can be found in specialist books on this topic, for example:
• Bannantine, J.A., Comer, J.J., Handrock, J.L., 1998, Fundamentals of
Metal Fatigue Analysis, Second Edition, Prentice Hall, Englewood
Cliffs, New Jersey.
• Schijve, J., 2009, Fatigue of Structures and materials, Second Edition,
Springer Netherlands, Dordrecht, the Netherlands.
Holistic structural integrity process
(HOLSIP) development
Holistic life (with all intrinsic/extrinsic factors)
P1:
Nucleation
As-manufactured, IDS
Crack/corrosion
Short cracks
LEFM
WFD/MSD
LEFM/EPFM
Repairable
... ...
Damage interaction
MSD interaction Fract. toughness
Residual strength
EPFM/LEFM
Special NDI
NDI detectable
/fretting nucleation
Noncontinuum mech.
Durability
Safe life (no env.) Damage tolerance (no env.)
P2: Short
crack
P3: Long
crack
P4:
Instability
Fig. 4 Overview of the HOLSIP programme [29]: IDS ¼ initial discontinuity state;
EPFM ¼ elastic plastic fracture mechanics; LEFM ¼ linear elastic fracture mechanics;
MSD ¼ multiple site fatigue damage; NDI ¼ nondestructive inspection;
WFD ¼ widespread fatigue damage.
27
Fatigue Requirements for Aircraft Structures](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-35-2048.jpg)
![Two especially relevant reports in the present context are:
• Hu,W.,Tong,Y.C.,Walker,K.F.,Mongru,D.,Amaratunga,R.,Jackson,
P.,2006,A review andassessmentofcurrent lifingmethodologiesandtools
in Air Vehicles Division, DSTO Research Report DSTO-RR-0321,
Defence Science and Technology Organisation, Melbourne, Australia.
• Molent, L., Barter, S.A., Wanhill, R.J.H., 2010, The lead crack fatigue
lifing framework, DSTO Research Report DSTO-RR-0353, Defence
Science and Technology Organisation, Melbourne, Australia.
4.1 Safe-Life Estimations
Stress-life approach: This is the traditional basis for Safe-Life fatigue analysis
of aircraft structures. A generic overview of the procedure is shown in Fig. 5.
Two points require elaboration:
(1) Coupon and component S–N curves are modified into the so-called
working curves with the aid of several adjustment factors that are usually
proprietary. These factors account for various influences on service
components that are not included or experienced by coupons and com-
ponents during the laboratory tests.
(2) The often-used Pålmgren-Miner cumulative damage rule ignores load
sequence effects, and the damage accumulation rate is assumed (incor-
rectly) to be independent of stress amplitude. The predicted lives can
range from 0.3 to at least 3 the actual life [30]. The potential for
unconservative life predictions (values 1) is mitigated by using scatter
factors based on engineering judgement and experience, as mentioned
in Section 3.
Structurally significant components
Component stress spectrum
Cumulative damage fatigue analysis
• Pålmgren-Miner rule to calculate life
• Retirement time ≤ safe life
Retirement / replacement
Structural usage fatigue loads
Coupon and component fatigue tests
• S–N tests S–N curves S–N
safe life
on calculated life
working curves
• Scatter factor
Fig. 5 Generic procedure for estimating the fatigue Safe-Life using the stress–life
approach.
28 Russell J.H. Wanhill](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-36-2048.jpg)
![Strain-Life approach: This is commonly used to analyse fatigue damage in
aircraft fleets. The damage (loads and accelerations) data are acquired via
strain recorders. There are several OEM and publicly available tools for
analysing the data.
Fig. 6 shows a generic strain–life procedure to estimate safe fatigue lives
and remaining lives from fleet data. In practice the test and service data and
cumulative damage analyses are integrated into a computer programme that
includes an algorithm for cycle counting. This is usually ‘Rainflow’ cou-
nting, which is well described in Bannantine et al. [31].
4.2 Damage Tolerance Estimations
DT approach: A generic analysis approach based on the USAF DT method is
shown in Fig. 7. As mentioned in Section 3, this method is used for many
USAF aircraft and has been effective in ensuring structural safety. There are
two points particularly to note:
(1) The FCG analyses use linear elastic fracture mechanics (LEFM) FCG
models. These models have difficulties in predicting the behaviour of
short fatigue cracks, i.e. cracks with dimensions less than about
0.5 mm [32,33].
(2) Fracture toughness data and residual strength analyses and tests are
required to complete the DT procedure.
Structurally significant components
· εa–N tests
εeq.a equation
Sl–εl
· Strain–life equation
· Notch analysis rule damage from εa–N equation
· Damage accumulation rule to calculate life
· Component stress spectra
· Component strain spectra
· Cycle counting (hysteresis loops)
Cumulative damage fatigue analysis
· S–ε cyclic stress–strain curve
Structural usage fatigue loads
· Scatter factor on calculated life
· Retirement time ≤safe life
Retirement / replacement
safe life
Fig. 6 Generic procedure for estimating the safe fatigue life and remaining fatigue life
using the strain–life approach.
29
Fatigue Requirements for Aircraft Structures](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-37-2048.jpg)
![DSTO approach: This differs importantly from the DT approach with
respect to the required input data and the FCG analysis methods, as may
be seen from Table 1. The following points bear repeating with the addition
of some important details:
(1) The DSTO FCG analyses begin with EPS representing actual initial dis-
continuity sizes (‘discontinuity’ is neutral and covers a variety of fatigue-
nucleating features; whereas ‘flaw’ is negative and can be inappropriate).
Use of actual initial discontinuity data was also the original intention
of the USAF DT approach for durability analyses, but until QF advances
since 1990 there was a lack of FCG data for (very) early crack growth
from initiation sites with EPS values less than about 0.25 mm.
Besides being important for durability analyses, service experi-
ence has shown the early FCG regime to contribute to improved
safety analyses for high-performance aircraft [10,34]. Also, the
importance of WFD (notably in large transport aircraft, see
Section 1.1) is another motivation for QF-based analyses of early
FCG and its development [35].
(2) The intransigence of the early FCG regime with respect to LEFM
modelling has shown the need to develop empirical relationships
describing the early FCG behaviour [10,34]. In turn, this requires
QF data compilations of early FCG from service aircraft, FSFT, and
component and coupon tests.
Hybrid approaches: Notwithstanding the differences between the USAF DT
and DSTO approaches, there is much to be said in favour of a rapprochement.
The DSTO’s (now DSTG) position is that both LEFM and empirical
FCG modelling can be used where justified by data obtained from service,
FSFT and component and coupon tests. Also, the EPS concept for early
crack growth has recently [36] been shown to be compatible with the USAF
EIFS durability concept mentioned in Section 3.
4.3 Operational Damage Estimations
Much effort is currently being put into structural health monitoring (SHM).
The evolution of SHM has followed two main strategies:
(1) Developing rapid NDI technologies to reduce the inspection time on
the ground, often obtaining help from on-board sensing systems.
(2) Monitoring of loads or strains or damage or a combination of these to
determine the remaining useful life (RUL) and/or setting up inspection
intervals for safe operation of the aircraft. Some of these systems have
evolved into health and usage monitoring systems (HUMS) which
use strain sensors extensively.
31
Fatigue Requirements for Aircraft Structures](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-39-2048.jpg)
![HUMS began and continues to be the basic system for helicopters, but it is
now being used on fixed-wing aircraft, especially tactical aircraft like the
Eurofighter and Indian Light Combat Aircraft (LCA). Strain-sensing equip-
ment installed on the aircraft provide operational data that are usually used to
estimate the RUL via a Safe-Life analysis, as shown generically in Fig. 6.
However, it is also possible to use a HUMS system in a DT analysis to
calculate a safe crack growth life [37–39]. One such system, developed by
the NLR and called ODAT (operational damage assessment tool), is shown
in Fig. 8. The architecture reflects the overall procedure for LEFM
Operational
Damage
Assessment
Output module
Admin module
MatDat module
K-fact module
CraGro module
Specturm module
Mission module
GeoMat module
login module h/c type version
External data
and tools
fleet data
representative
mission profiles
trained
models
NLR
ANN tool
all
mission
profiles
tail number
main assembly
h/c
reference
data
meas.
missions
subset
ANN
models
HELIUM
sub assembly
crack characteristics
component material type
user rights module
activity log module
dbase mgmt module
material data
K-solutions
crack size
stress scale factor
stress
sequences
sequence
generator
crit
det
flhrs
da/dN = f(ΔK,R)
K = f(crack size)
projected missions
available missions
mission mix
flight parameter
sequences
Remaining flight hours OK?
yes no
Tool
Fig. 8 The ODAT system architecture developed by the NLR for a helicopter fa-
tigue life monitoring programme: HELIUM ¼ helicopter usage monitoring database;
ANN ¼ artificial neural network. (Illustration courtesy of Marcel Bos, NLR.)
32 Russell J.H. Wanhill](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-40-2048.jpg)
![(2) Even when the most critical feature of a component has been identified
and assessed by coupon testing, the coupons are rarely fully representa-
tive, notably with respect to the surface treatments and finishes required
for production aircraft. This is important because the commencement
of fatigue cracking is primarily surface-influenced and therefore greatly
dependent on small surface discontinuities inherent to component pro-
duction, as well as any surface-connected discontinuities inherent to the
material.
These limitations are addressed by other means. One way, which is mandatory
for all modern aircraft, is to test actual components, and conduct FSFT on part
of the structure or even the full airframe, thereby including the effects of com-
ponent geometry and production. (N.B.: FSFT became mandatory for mil-
itary aircraft in 1969, and civil aircraft in 1998, see Sections 1.1 and 1.2)
Another way is to improve coupon testing by making the coupons opti-
mally representative of the most fatigue-critical details, e.g. by applying sur-
face treatments and finishes used in component production. This may seem
obvious, but it is sometimes neglected or overlooked.
5.1 ‘Building Block’ Testing Procedure
Fig. 9 shows a schematic ‘Building Block’ (BB) approach for testing mate-
rials, components and structures as part of an aircraft certification process. This is
Ground flight tests
Components, FSFT
Subcomponents
Elements
Coupons
P
r
e
-
d
e
s
i
g
n
t
e
s
t
s
P
o
s
t
-
d
e
s
i
g
n
t
e
s
t
s
Fig. 9 ‘Building Block’ fatigue test approach for materials, components and structures:
after [40].
34 Russell J.H. Wanhill](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-42-2048.jpg)
![adapted from a schematic for the Lockheed Martin F-35 Joint Strike Fighter
(JSF), but is generically valid. The BB approach may be viewed as a pyramid
whose base is the initial material evaluation. Each level of the pyramid is the
foundation for the next, and the structural complexity and costs increase
with each level up to the FSFT(s). The final phase of certification is ground
and flight testing of the aircraft.
There are several points to be noted about the testing approach and
procedure:
Coupon testing level: This includes standard and nonstandard tests.
Examples of standard tests are stress–life (S–N); strain–life (ε–N); cyclic
stress–strain (S–ε); and LEFM FCG. The latter are used to estimate FCG
lives according to the USAF DT fatigue lifing method, see Fig. 7. An appro-
priate source for standard tests is ASTM International.
Nonstandard coupon tests have been developed for the DSTO approach
to DT analysis. One type of coupon is a simple flat specimen with an array of
laser-induced crack starter slots. These enable QF measurements of FCG
from initial discontinuities 0.05 mm in size.
However, QF usually requires ‘marker loads’ to be inserted into the
fatigue load histories. The types of markers, and the ways in which they
are added, can be chosen to ensure negligible influences of the marker loads
on the overall FCG [41].
Element to FSFT levels: At all these levels it is advisable to add marker
loads to the fatigue load histories. Sometimes a realistic load history will
result in natural crack front markers, but it is better to make sure that they
occur. There are comprehensive guidelines for this [41].
Crack front marking is especially relevant to QF analyses of fatigue cracks
detected during FSFT teardowns, see Fig. 10. Besides being used (and
required) directly for certification, the FSFT teardown results are important
for possible design modifications, verifying structural analyses, and deter-
mining whether retrofits may be advisable or necessary.
FSFT types: It is important to note that FSFT does not imply testing of
the entire airframe. This may be done in some instances, but the available
space and testing equipment often dictate that major parts of the airframe
are tested separately.
An example of part-structure FSFT is the Airbus A380 MegaLiner Barrel
(MLB) fatigue test, in which candidate fuselage skin materials were tested.
Fig. 11 shows (i) the MLB; (ii) the types of applied loads, namely fuselage
pressurisation (ΔP) and bending (MY, MZ) and ground loads (QZ); and
(iii) the number of simulated flights applied during testing. This is slightly
more than twice the nominal Design Service Goal (DSG) of 20,000 flights.
35
Fatigue Requirements for Aircraft Structures](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-43-2048.jpg)
![FSFT requirements: As mentioned in Section 1.1, cases of WFD in civil
transport aircraft led eventually to mandatory FSFT from 1998 onwards; and
in 2011 the LOV concept for aircraft above 34,000 kg. The LOV concept
requires FSFT to determine the onset of WFD [11].
These changes were reflected in the FSFT requirements, or rather expec-
tations. From 1998 to 2011 the expectation was that an FSFT would be done
Fig. 10 Airframe fatigue certification and the central role of fatigue testing: see Fig. 9
also: original illustration courtesy of Madeleine Burchill and Simon Barter, DSTG,
Melbourne.
MZ
MZ
QZ
ΔP
45,402 simulated flights with
conservatively high fatigue loads
MY
MY
Fig. 11 Airbus A380 MegaLiner Barrel (MLB) full-scale fatigue test ‘specimen’, the gen-
eral loading conditions, and the number of simulated flights applied during testing [42].
36 Russell J.H. Wanhill](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-44-2048.jpg)
![to a minimum of two DSGs followed by specific inspections and analyses [8].
With introduction of the LOV concept it is recommended that the FSFT is
run to three DSGs, followed by residual strength testing [8].
The USAF has followed a similar evolutionary path in its DT approach.
Thus MIL-STD-1530C [43], which predates the LOV concept, suggests
that the FSFT be extended beyond two DSGs because the onset of WFD
does not always occur within the two-lifetime test period [6].
6 SUMMARY
The fatigue requirements for aircraft structures have evolved since the
1950s. However, this has not been a gradual process: ‘milestone’ accidents
have caused paradigm shifts for both civil and military aircraft, as may be seen
from Figs 1 and 2.
The requirements encompass fatigue design and analysis, and verification
using several levels of testing. The ‘final’ verification is by FSFT and tear-
downs, which provide checks on designs and analyses, and can reveal the
need for modifications and retrofits.
Besides ‘final’ verification of structural integrity, FSFT and teardowns
provide references for any fatigue problems that arise during service. They
are also references for investigating whether it is feasible to extend the service
life beyond the original design life. This is a not infrequent issue.
ACKNOWLEDGEMENTS
Thanks are due to Simon Barter, Lorrie Molent and Madeleine Burchill, DSTG, Melbourne,
Australia; and Marcel Bos, NLR, Emmeloord, the Netherlands.
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![Application of Titanium Alloys in Military Platforms, RTO Meeting Proceedings 69
(II), NATO Research and Technology Organisation, Neuilly-sur-Seine, France,
2003, pp. 1-1–1-11.
[23] R.J.H. Wanhill, Aerospace applications of aluminum-lithium alloys, in: N.E. Prasad, A.
A. Gokhale, R.J.H. Wanhill (Eds.), Aluminum-Lithium Alloys, Processing, Properties
and Applications, Butterworth-Heinemann, Elsevier Inc., Oxford, UK, 2013,
pp. 503–535.
[24] Beumler, T., Flying GLARE®
, a contribution to aircraft certification issues in non-
damaged and fatigue damaged GLARE®
structures, (Doctor’s Thesis), Delft University
of Technology, Delft, The Netherlands, 2004.
[25] Federal Aviation Administration, Composite Aircraft Structure, Advisory Circular FAA
AC-20-107B, Change 1, August 24, 2010, U.S. Department of Transportation,
Washington, DC, 2010.
[26] MIL-HDBK-17-3F, Military composite materials handbook, in: Polymer Matrix
Composites Materials Usage, Design, and Analysis, vol. 3, U.S. Department of Defense,
The Pentagon, VA, 2013.
[27] Ilcewicz, L., 2007, AC 20-107A “Composite Aircraft Structure” updates, Commercial
Aircraft Composite Repair Committee Meeting, Wichita, Kansas, 14 November, 2007.
[28] Ilcewicz, L., 2009, Updates to AC 20-107B “Composite Aircraft Structure”, Compos-
ite Damage Tolerance Maintenance Workshop, Tokyo, Japan, 5 June 2009.
[29] J.P. Komorowski, N.C. Bellinger, M. Liao, A. Fillion, Application of the holistic struc-
tural integrity process to Canadian Forces challenges, in: ASIP 2007, USAF Aircraft
Structural Integrity Program 2007, December 4–6, 2007, Palm Springs, CA, 2007.
[30] W. Sch€
utz, The prediction of fatigue life in the crack initiation and propagation stages—
a state of the art survey, Eng. Fract. Mech. 11 (1979) 405–421.
[31] J.A. Bannantine, J.J. Comer, J.L. Handrock, Fundamentals of Metal Fatigue Analysis,
second ed., Prentice Hall, Englewood Cliffs, NJ, 1998.
[32] R.F.W. Anstee, An assessment of the importance of small crack growth to aircraft
design, in: Behaviour of Short Cracks in Aircraft Components, AGARD Conference
Proceedings No. 328, Advisory Group for Aerospace Research and Development,
Neuilly-sur-Seine, France, 1983, pp. 3-1–3-9.
[33] R.F.W. Anstee, P.R. Edwards, A review of crack growth threshold and crack
propagation rates at short crack lengths, in: Some Considerations on Short Crack
Growth Behaviour in Aircraft Structures, AGARD Report No. 696, Advisory Group
for Aerospace Research and Development, Neuilly-sur-Seine, France, 1983,
pp. 2-1–2-12.
[34] L. Molent, S.A. Barter, R.J.H. Wanhill, The Lead Crack Fatigue Lifing Framework,
DSTO Research Report DSTO-RR-0353Defence Science and Technology Organi-
sation, Melbourne, Australia, 2010.
[35] R.J.H. Wanhill, M.F.J. Koolloos, Fatigue and corrosion in aircraft pressure cabin lap
splices, Int. J. Fatigue 23 (2001) S337–S347.
[36] J. Gallagher, L. Molent, The equivalence of EPS and EIFS based on the same crack
growth life data, Int. J. Fatigue 80 (2015) 162–170.
[37] J.B. de Jonge, C.W.J. van Lummel, Development of a Crack Severity Index (CSI) for
Quantification of Recorded Stress Spectra, NLR Technical Report NLR-TR-86049,
National Aerospace Laboratory NLR, Amsterdam, The Netherlands, 1986.
[38] H. Lee, S. Park, H. Kim, Estimation of aircraft structural fatigue life using the crack
severity index methodology, J. Aircr. 47 (5) (2010) 1672–1678.
[39] Aeronautical Systems Center Engineering Directorate, Methodology for Determina-
tion of Equivalent Flight Hours and Approaches to Communicate Usage Severity,
Structures Bulletin EN-SB-09-001, ASC/EN, Bldg 560, 2530 Loop Road West,
Wright-Patterson AFB, OH, 2009.
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![[40] D.L. Ball, D.S. Norwood, S.C. TerMaath, Joint Strike Fighter airframe durability and
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[41] S.A. Barter, R.J.H. Wanhill, Marker Loads for Quantitative Fractography (QF) of
Fatigue in Aerospace Alloys, NLR Technical Report NLR-TR-2008-644, National
Aerospace Laboratory NLR, Amsterdam, the Netherlands, 2008.
[42] R.J.H. Wanhill, D.J. Platenkamp, T. Hattenberg, A.F. Bosch, P.H. de Haan, GLARE
teardowns from the MegaLiner Barrel (MLB) fatigue test, in: M.J. Bos (Ed.), ICAF
2009: Bridging the Gap between Theory and Operational Practice, Springer,
Netherlands, Dordrecht, the Netherlands, 2009, pp. 145–167.
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40 Russell J.H. Wanhill](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-48-2048.jpg)
![CHAPTER THREE
Typical Fatigue-Nucleating
Discontinuities in Metallic
Aircraft Structures
Simon A. Barter*, Lorrie Molent*, Russell J.H. Wanhill†
*Aerospace Division, Defence Science and Technology Group, Australia
†
Aerospace Vehicles Division, National Aerospace Laboratory NLR, The Netherlands
1 INTRODUCTION
Under more or less ideal conditions (highly polished surfaces in
homogeneous materials), a significant period of fatigue-induced microstruc-
tural damage can precede fatigue crack nucleation. However, production
aircraft components and structures have many sources of surface or near-
surface discontinuities from which fatigue cracking may very quickly
develop in highly stressed locations once loading has commenced, i.e. upon
entry of the aircraft into service.
The discontinuities include various forms of machining damage
(scratches, grooves, burrs, small tears and nicks); etch pits from surface treat-
ments (pickling, anodising); porosity, especially in thick aluminium alloy
plate and castings; and in the case of aluminium alloys and steels, constituent
particles that may themselves be cracked.
A fatigue lifing framework using a lead crack concept1
has been devel-
oped by the DST Group for metallic primary airframe components [1,2].
Like the original damage tolerance (DT) concept developed by the United
States Air Force (USAF), this framework assumes that fatigue cracking
begins as soon as an aircraft enters service. However, there are major and
fundamental differences. Instead of assuming initial crack sizes and deriving
early crack growth behaviour from back-extrapolation of growth data for
long cracks, the DST framework uses data for real cracks growing from small
discontinuities inherent to the materials and component production. To this
end, this paper examines the types of discontinuities that act as a nucleus for
fatigue cracks in typical metallic airframe structures with a view of informing
1
Lead cracks are fatigue cracks that grow faster (or fastest) in a particular structural area.
Aircraft Sustainment and Repair © 2018 Elsevier Ltd.
https://doi.org/10.1016/B978-0-08-100540-8.00003-0 All rights reserved.
41](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-49-2048.jpg)
![preservice NDI. Other types of discontinuity are simply unavoidable owing
to limitations of the design and production processes, e.g. etch pits from
chemical treatments, porosity in castings (which have only limited use in
primary airframe structures), and the constituent particles in aluminium
alloys and steels.
In all these cases, the presence and abundance of the discontinuities,
especially combined with high local stresses, result in early fatigue cracking
that leads any other fatigue cracking that may occur.
The main objectives of this paper are (a) to present examples of the types
of discontinuities that actually nucleate fatigue cracks in typical metallic
airframe structures and (b) enable designers, production and maintenance
organisations to more fully understand the source and nature of what is
usually unexpected fatigue cracking. The examples are for cracks from
coupon, component and full-scale tests, and also from service aircraft, includ-
ing commercial transport aircraft and high performance military aircraft.
In every one of the presented examples, the time to fatigue crack
nucleatation was found by nondestructive inspection (NDI) or quantitative
fractography (QF) to be very short: in fact, so short as to permit the assump-
tion of immediate crack growth for lead cracks when applying the lead crack
concept [1,2].
2 EXAMPLES OF DISCONTINUITIES
2.1 Poorly Finished Holes
Fastener hole drill or reamer damage, burrs, and poor de-burring practice are
common production defects encountered in metallic airframes. Table 2 lists
the examples considered here and their sources: Examples 1 and 3–5 are
from Aermacchi MB326 trainers, now retired from RAAF service;
Example 2 is from a FSFT of the Airbus MegaLiner Barrel (MLB);
Example 6 is from a Dassault Mirage III, also withdrawn from RAAF ser-
vice; Example 7 is from coupon tests used to support the Boeing F/A-18A/
B lifing programme; Examples 8 and 9 are from a FSFT of an ex-service
Lockheed P3C empennage.
Example 1. Fig. 1 shows the fatigue cracking, which started on both sides of
the fastener hole at the beginning of the FSFT and ultimately failed the wing
main spar. On the right-hand side of the hole (as shown in the figure) three
cracks (arrowed) grew from gross drill-damage. These cracks interacted
and coalesced, and hence grew faster than a single crack growing from a
43
Typical Fatigue-Nucleating Discontinuities in Metallic Aircraft Structures](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-51-2048.jpg)
![drill-induced tear on the left-hand side. The red arrow on this side indicates
the extent of this crack when the right-hand side failed. Quantitative
fractography (QF) of the right-hand side cracks showed that they most
probably started during the application of the first block of FSFT loading.
Table 2 Examples of Fatigue-Nucleating Discontinuities: Poorly Finished Holes
Discontinuity
Description Component Aircraft Type Source No. Ref.
Drilling: gross damage,
tearing
Drilling: scoring
Fastener insertion with
scoring
Bad de-burring
Machining tears
AA7075-T6 main spar
GLAREa
door beam
AA7075-T6 main spar
AA7075-T6 main spar
AISI 4340 carry-
through beam
AU4SGb
spar
AA7050-T7451
coupon
AA7075-T6
empennage
Aermacchi
MB326
Airbus MLB
Aermacchi
MB326
Aermacchi
MB326
Aermacchi
MB326
Dassault
Mirage III
Boeing F/A-
18A/B
Lockheed
P3C
FSFT
FSFT
Service
Service
Service
Service
Test
FSFT
1
2
3
4
5
6
7
8
[3]
[4]
[3]
[3]
[5]
[6]
[7]
[8]
a
GLARE ¼ glass reinforced aluminium laminates.
b
AU4SG AA2014-T6.
1 mm
Fig. 1 Example of a poorly drilled hole in a main wing spar made from an AA7075-T6
aluminium alloy extrusion. The cracking on the right side of the hole has nucleated
from gross drill damage, while the crack on the left side of the hole nucleated from
a drill-induced surface tear during machining. The arrow on the far left indicates the
position of the fatigue crack on this side when the right ligament failed.
44 Simon A. Barter et al.](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-52-2048.jpg)
![Example 2. Fig. 2 shows two linked-up FSFT fatigue cracks in the bore of a
GLARE fastener hole in the MLB passenger door beam. Most cracks grew
from the 2024-T3 aluminium alloy sheet corners [4], but in this example a
crack also started from scoring owing to poor drilling. This result and other
evidence of poor drilling led to a production improvement programme for
GLARE components in the Airbus A380. Again, posttest QF indicated that
the cracks most probably started very early in the FSFT.
Example 3. Fig. 3 shows service fatigue cracks starting from scoring along
the bore of a fastener hole in a wing main spar. The scoring was due to inser-
tion of a blind fastener. The cracks grew more or less evenly on both sides
of the hole, and would probably have failed the spar if it had remained in
service. Discovery of these cracks led to an extensive teardown programme,
and many similarly damaged holes were found [3]. An analysis of fatigue
metre data and a correlation of these data with the progression markings
on the fatigue fracture surfaces indicated that the cracking had started very
early in the life of the aircraft.
Example 4. Fig. 4 shows service fatigue cracks growing from a fastener hole
in a wing main spar withdrawn from service. The hole had been badly
de-burred, resulting in a split burr that started the upper crack and several
nicks (arrowed), one of which started the lower crack. These cracks were
found during the teardown programme mentioned for Example 3 [3].
200µm
Fig. 2 Example of fatigue cracking nucleation from scoring in the bore of a fastener hole
in the MLB GLARE door beam [4]. The cracking is in one of the AA2024-T3 sheets of
the GLARE laminate.
45
Typical Fatigue-Nucleating Discontinuities in Metallic Aircraft Structures](https://image.slidesharecdn.com/23391083-250523091517-ccee1d2e/75/Aircraft-Sustainment-And-Repair-Rhys-Jones-Alan-Baker-Neil-Matthews-53-2048.jpg)
Aircraft Sustainment And Repair Rhys Jones Alan Baker Neil Matthews
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- 6. AIRCRAFT SUSTAINMENT AND REPAIR Editedby RHYS JONES ALAN BAKER NEIL MATTHEWS VICTOR CHAMPAGNE
- 7. Butterworth-Heinemann is animprint of Elsevier The Boulevard, Langford Lane, Kidlington, Oxford OX5 1GB, United Kingdom 50 Hampshire Street, 5th Floor, Cambridge, MA 02139, United States © 2018 Elsevier Ltd. All rights reserved. Victor Champagne is an US Gov employee and his contributions are in public domain. No part of this publication may be reproduced or transmitted in any form or by any means, electronic or mechanical, including photocopying, recording, or any information storage and retrieval system, without permission in writing from the publisher. Details on how to seek permission, further information about the Publisher’s permissions policies and our arrangements with organizations such as the Copyright Clearance Center and the Copyright Licensing Agency, can be found at our website: www.elsevier.com/permissions. This book and the individual contributions contained in it are protected under copyright by the Publisher (other than as may be noted herein). Notices Knowledge and best practice in this field are constantly changing. As new research and experience broaden our understanding, changes in research methods, professional practices, or medical treatment may become necessary. Practitioners and researchers must always rely on their own experience and knowledge in evaluating and using any information, methods, compounds, or experiments described herein. In using such information or methods they should be mindful of their own safety and the safety of others, including parties for whom they have a professional responsibility. To the fullest extent of the law, neither the Publisher nor the authors, contributors, or editors, assume any liability for any injury and/or damage to persons or property as a matter of products liability, negligence or otherwise, or from any use or operation of any methods, products, instructions, or ideas contained in the material herein. Library of Congress Cataloging-in-Publication Data A catalog record for this book is available from the Library of Congress British Library Cataloguing-in-Publication Data A catalogue record for this book is available from the British Library ISBN: 978-0-08-100540-8 For information on all Butterworth-Heinemann publications visit our website at https://www.elsevier.com/books-and-journals Publisher: Matthew Deans Acquisition Editor: Christina Gifford Editorial Project Manager: Jennifer Pierce Production Project Manager: Vijayaraj Purushothaman Cover Designer: Victoria Pearson Typeset by SPi Global, India
- 8. CONTRIBUTORS David R. Arnott DefenceScience & Technology Group, Fishermans Bend, VIC, Australia Alan A. Baker Advanced Composite Structures Australia; Defence Science and Technology Group, Fairbairn, CBR, Australia; Aerospace Division, Defence Science and Technology Group, Melbourne, VIC, Australia Simon A. Barter Aerospace Division, Defence Science and Technology Group, Australia Simon Barter Defence Science and Technology Organisation, Fishermans Bend, VIC, Australia Andreas J. Brunner Empa, Swiss Federal Laboratories for Materials Science and Technology, D€ ubendorf, Switzerland Victor Champagne US Army Research Laboratory, Aberdeen, MD; University of Massachusetts, Amherst, MA, United States Wenchen Hu Monash University, Clayton, VIC, Australia Pu Huang Monash University, Clayton, VIC, Australia David Hui University of New Orleans, New Orleans, LA, United States Rhys Jones Monash University, Clayton, VIC, Australia; Wrexham Glyndwr University, Wrexham, United Kingdom Anthony J. Kinloch Imperial College London, London, United Kingdom Neil Matthews RUAG Australia, Bayswater, VIC, Australia James J. Mazza U.S. Air Force Research Laboratory (AFRL/RXSA), Dayton, OH, United States Alison J. McMillan Wrexham Glyndwr University, Wrexham, United Kingdom John G. Michopoulos US Naval Research Laboratory, Washington, DC, United States xiii
- 9. Lorrie Molent Aerospace Division,Defence Science and Technology Group, Australia; Defence Science and Technology Group, Fishermans Bend, VIC, Australia Trung Nguyen Naval Air Systems Command, Patuxent River, MD, United States Daren Peng Monash University, Clayton, VIC, Australia Nam Phan Naval Air Systems Command, Patuxent River, MD, United States Kirk Rackow Sandia National Laboratories, Albuquerque, NM, United States Andrew N. Rider Defence Science & Technology Group, Fishermans Bend, VIC, Australia Dennis Roach Sandia National Laboratories, Albuquerque, NM, United States John Wang Aerospace Division, Defence Science and Technology Group, Melbourne, VIC, Australia Russell J.H. Wanhill Emmeloord; Aerospace Vehicles Division, National Aerospace Laboratory NLR, The Netherlands xiv Contributors
- 10. CHAPTER ONE Rationale andOrganisation Rhys Jones*, Neil Matthews† , Alan A. Baker‡,§ , Victor Champagne Jr.¶ *Monash University, Clayton, VIC, Australia † RUAG Australia, Bayswater, VIC, Australia ‡ Advanced Composite Structures Australia, Melbourne, VIC, Australia § Defence Science and Technology Group, Fairbairn, CBR, Australia ¶ US Army Research Laboratory, Aberdeen, MD, United States In general, the design of aerospace vehicles requires that all structures and repairs to load-bearing structural elements be designed in accordance with damage tolerance design principles, which are largely based on the discipline of fracture mechanics and which for military aircraft are detailed in the Joint Services Structural Guidelines-2006 (JSSG2006) [1] and in the USAF Dam- age Tolerant Design Handbook [2]. This design philosophy has evolved as a result of a number of high-profile (military aircraft) incidents some of which,1 together with a number of civilian aircraft failures, are reviewed in Ref. [3]. As explained in Ref. [3]: While there have been lessons learned from nearly all of the accidents and inci- dents, those that were seminal with respect to the design and certification of air- craft structures were the 1954 Comet failures, the 1958 B-47 accidents, the 1969 F-111 accident, the 1976 AVRO 748 accident, the 1978 Dan Air 707 Lusaka accident and the 1988 Aloha Boeing 737 accident. The Comet failures resulted in the use of fail-safe design in commercial aircraft; the B-47 accidents resulted in the development of the Air Force Aircraft Structural Integrity Programme (ASIP); the F-111 accident accelerated the introduction of damage tolerance design requirements in military aircraft; the AVRO 748 and Dan Air Boeing 707 accidents resulted in complementing commercial fail-safe design with damage tolerance requirements; and the Aloha Boeing 737 accident brought attention to the impor- tance that WFD2 has in limiting the safe operation of airplane structures. The USAF [5] was one of the first to realise that the operational life of an aircraft was governed by the growth of small cracks and that the analysis tools 1 Ref. [3] also provides a historical perspective as to how the military, the Federal Aviation Administration (FAA) and industry have dealt with these threats. 2 Widespread fatigue damage (WFD) is defined in Ref. [4] as the simultaneous presence of cracks at multiple structural locations that are of sufficient size and density such that the structure will no longer meet the residual strength requirements of FAR section §25.571(b). Aircraft Sustainment and Repair © 2018 Elsevier Ltd. https://doi.org/10.1016/B978-0-08-100540-8.00001-7 All rights reserved. 1
- 11. needed for aircraftsustainment issues, that is, to assess in-service cracking, differed from those used in design. This conclusion was also one of the pri- mary conclusions reached in the USAF-McDonnell Douglas (now Boeing) study into sustainment issues associated with cracking in the F-15 aircraft [6] and is also enunciated in Refs. [7,8]. It also led to the USAF developing a probabilistic approach to assessing the fatigue life of operational aircraft [5] in which crack growth was represented as an exponential function, viz.: a ¼ aieωN (1) Here a is the crack length/depth; ai is the initial crack size, which is often referred to as the equivalent initial flaw size (EIFS) or the equivalent precrack size (EPS) [9]; N is the number of cycles, or flight hours; and ω is a constant. While the operators of military aircraft had long been aware of the problem of aging of an aircraft [3], public awareness of the problem first arose via the Aloha Airlines accident in April 1989 [10] which involved failure due to the presence of multisite damage in the fuselage lap joint of a Boeing 737-200 aircraft. Indeed, despite the fact that there were few casualties involved, it continues to be described as the event which first brought the aging aircraft issue to the attention of the general public and to those who have responsibility to ensure aircraft safety. The extent of the problems associated with fuselage lap joints in civil transport aircraft was further illustrated by the April 2011 incident whereby cracking in a fuselage lap joint in a Southwest Airlines Boeing 737-300 aircraft resulted in a large 1.52 m (5 ft) hole in the roof [11]. This incident led to the grounding of 79 of Southwest’s older Boeing 737 aircraft [11] and to the cancellation of almost 700 flights. Subsequent inspections, which found cracks in a total of four Southwest aircraft [12], led to the US Federal Aviation Adminis- tration (FAA) mandating the inspection of 175 Boeing 737 aircraft that had seen more than 35,000 pressurisation cycles. The problem of cracking in fuselage lap joints was not confined to Boeing 737 aircraft. On 26 Octo- ber 2010, an American Airlines 757-200 aircraft was forced to land at Miami International Airport due to a sudden decompression arising from cracking in a fuselage joint [13]. This aircraft had experienced less than 23,000 cycles. This led to the discovery of cracking in other 757 aircraft and a subsequent January 2011 FAA Airworthiness Directive [13] mandat- ing the inspection of all 757-200 and 757-300 aircraft. 2 Rhys Jones et al.
- 12. As a resultof these incidents, and other considerations, the FAA intro- duced the concept of a limit of viability (LOV), defined as the onset of multisite and/or multielement damage, which the FAA now uses to define (or limit) the operational life of civil transport aircraft [13,14]. Thus given the central role that fatigue, fracture mechanics and damage tolerance play in aircraft design and sustainment, the first four chapters of this book focus on the following: • The evolution of fatigue design requirements for aircraft structures with an emphasis on requirements, with design and operational issues also being addressed (Chapter Two). • Typical fatigue initiating discontinuities in metallic aircraft structures and the lead crack concept. (Chapter Three). • Practical computational fracture mechanics for use in assessing aircraft structural integrity (Chapter Four). • Methods for computing/predicting the growth of cracks that arise from naturally occurring material discontinuities in operational aircraft (Chapter Five). Chapters Two–Five are closely linked in that they each discuss the fact that the growth of ‘lead’ cracks [15], which are defined as the fastest growing cracks in the airframe and which (as a result) set both the inspection intervals and the operational life of the fleet, is generally exponential and as such fol- lows Eq. (1). Chapter Two shows that, in the absence of severe corrosion, for aluminium alloy airframes the mean value of the EPS (EIFS) is 0.01 mm. Chapter Five shows that the exponent ω in the exponential growth law is approximately proportional to the cube of the stress [16]. Chapter Five explains that this cubic approximation is now used by the Royal Australian Air Force (RAAF) for the F/A-18 Classic Hornet, the AP3C (Orion) and the Pilatus Porter PC9 fleets [17–19]. It is further shown that crack growth in operational aircraft and in full-scale fatigue tests tested under an operational flight load spectra can be collapsed onto a single ‘master curve’ regardless of the material or the flight load spectrum. Chapter Five also shows that, as shown in Ref. [8], the growth of both long and short cracks can be captured by the Hartman–Schijve variant of the NASGRO crack growth equation. In this context, it is shown that for cracks that arise and grow from small material discontinuities in operational aircraft the Hartman–Schijve equation can be approximated by a simple Paris crack growth equation with the exponent and the constant of propor- tionality determined from tests on ASTM E647-13a long crack specimens. 3 Rationale and Organisation
- 13. The ability ofthe Hartman–Schijve crack growth equation to compute accurately the growth of cracks in 7075-T6 P3C Orion test specimen with intergranular cracking at a dome nut hole that has been repaired using super- sonic particle deposition (SPD) is further highlighted in Chapter Sixteen. These findings are important since Chapters Twelve and Thirteen subse- quently show that a variant of the Hartman–Schijve equation with ΔK rep- laced by Δ√G, where G is the energy release rate, also captures the growth of disbonds in adhesively bonded joints and delamination growth in com- posites [20–23]. As such the Hartman–Schijve equation can be thought of as a unified formulation that can be used when assessing the original airframe and the damage tolerance of any subsequent structural repair. In this context, it should be noted that before the Hartman–Schijve equa- tion, or any other crack growth equation, can be used to predict/compute crack growth in an operational structure it is first necessary to determine the relationship between the stress intensity factor K and the crack length/ depth. To this end, Chapter Four presents simple computational tools, both three-dimensional (3D) weight function and 3D finite element-based approaches, for determining the stress intensity factors associated with small sub-mm cracks in realistic complex geometries under arbitrary flight loads. Since, as explained in the fatigue test standard ASTM E647-13a, the majority of the fatigue life of an airframe is generally consumed in growing from a small sub-mmmaterialdiscontinuitytoasizethatisreadilydetectableusingmodern nondestructive inspection techniques attention is focused on small initial cracks that first grow as a complex 3D shape and then transition into a through-the-thickness flaws with an oblique (part) elliptical crack front. At this stage, it should be noted that Kinzie and Peeler [24] have shown how the problem of aging aircraft is particularly acute in the military sphere and how it has been compounded by US Congress, Public Law 107-314 Sec: 1067 which is entitled: “Prevention and mitigation of corrosion of military equipment and infrastructure.” One impact of this law is that the effect of corrosion on the structural integrity of USAF, US Army, US Navy aircraft and other Navy weapons platforms must be now quantified and managed as per cracking, that is, in a damage-tolerant manner. In this con- text, the June 2007 Report to Congress by the Under Secretary of the Department of Defense (Acquisition, Technology and Logistics) [25] esti- mated the cost of corrosion associated with US Department of Defense (DoD) systems to be between $10 billion and $20 billion per year. This report also outlined the need for research into four primary areas. One of 4 Rhys Jones et al.
- 14. these research areasis: Repair processes that restore materials to an accept- able level of structural integrity and functionality. The subsequent chapters address methods that can be used to meet this challenge, viz.: externally bonded composite repairs, SPD, also known as cold spray (CS) [26], laser additive deposition (LAD) and multiplicative manufacturing, which involves the use of both subtractive and additive manufacturing. In this context, Chapters Six–Thirteen focus on composite repair tech- nology. This technology has been shown [27–33] to be highly successful in extending the fatigue lives of both primary and secondary (thin skin) struc- tures and in ensuring the continued safe operation of aging aircraft. Chapter Six highlights the key issues associated with bonded composite repairs and suggests approaches to repairs and reinforcement that contribute to extending the lifespan and/or the inspection interval when applied to pri- mary airframe structures. The focus is on the development of alternative approaches to conventional NDI to assess the structural integrity of the adhesive bond and based on these approaches the through-life management of composite repairs. Chapter Seven focuses on surface treatment and repair bonding and, as such, complements the work presented in Chapter Six. Both fundamental and practical aspects of bonded repair application are described. Chapter Seven also stresses the need for the reproducible durable bonds which is a fundamental requirement for bonded repair technology. Chapter Eight details the structural mechanics underpinning composite repairs to cracks in operational aircraft. Particular focus is given to: the role of adhesive and matrix nonlinearities; the importance of first ply failure in design and sustainment; application of the cubic rule in designing composite repairs and on assessing their effect on operational aircraft; and to the nature of the test specimens needed for certification. Chapter Eight also notes that, as explained in the seminal paper by Hart- Smith [34], adhesive performance is uniquely characterised by the strain energy density in the adhesive. As such, it clarifies one of the primary design methodologies enunciated in the PABST report [35], viz. that a critical design requirement is that the adhesive strain energy density should be kept beneath its fatigue threshold Wf. A simple means for obtaining Wf, which is based on the ASTM D1002 thick adherend test, is also presented. Chapter Eight also shows that for cracks that have grown from small nat- urally occurring material discontinuities, the role of the fibres bridging the crack is often a second-order effect. As a result, it is shown that the effect of a composite patch on crack growth in operational aircraft can be estimated 5 Rationale and Organisation
- 15. using the exponentialcrack growth approach presented in Chapter Five, and in the USAF approach to determine the risk of failure [5], together with the cubic rule, as also enunciated in Chapter Five, and the assumption that the effect of the patch is merely to lower the stress field in the underlying struc- ture. This finding is illustrated both via laboratory coupon tests and by com- parison with the test programme used to substantiate the boron epoxy repair of the USAF C-141 aircraft [36] and the repair to the F-111 lower wing skin [37]. Chapter Nine illustrates a building block approach, which uses the design methodology outlined in Chapter Eight, to outline the development of a composite repair to multisite damage in fuselage lap joints similar to those used in Boeing 737 and Airbus A320 civil transport aircraft. In this approach, the damage tolerance of the repair is first validated in the labora- tory and then confirmed by a series of flight demonstrator programmes. For the repair to the Airbus A320 fuselage lap joint, the damage tolerance of the repair was validated by application to a full-scale Airbus A320 fatigue test article. In one demonstrator programme, which was performed as part of the FAA Aging Aircraft programme [38], nine doublers were applied to a QANTAS B747-300 aircraft. These doublers experienced a variety of envi- ronmental conditions ranging from the tropics to Europe in mid-winter. These doublers were flown for 9 years and saw in excess of 37,000 h of service and 7020 landings. There was no disbonding or delamination asso- ciated with these doublers. Chapter Ten discusses the development of a bonded composite rein- forcement to the integrally stiffened D6ac steel wing pivot fitting (WPF) of F-111 aircraft in service with the RAAF. The high load transfer require- ments made this a very challenging application, especially as it was developed at an early stage in the development of this repair technology. Particular attention is again given to a building block approach and as such this chapter describes the finite element and the experimental programme used to deter- mine improved stiffener runout geometry as well as the full-scale wing tests used to validate the computed stress reduction. This repair represents the first bonded reinforcement to primary structure and was successful in that by reducing the strains at the critical location by 30%, it achieved its primary objective namely to stop cracking and wing failure in the RAAF F-111 fleet and cold proof load test (CPLT). Not only was the WPF reinforcement the first example of the ability of composite repair technology to eliminate cracking in primary structure, it was also the first example to highlight the benefit to be gained by combining composite repair technology with 6 Rhys Jones et al.
- 16. shape optimisation (reworkingthe local geometry). This approach subse- quently led to the development of a multiplicative manufacturing solution to cracking in the LAU7 missile launcher housing (see Chapter Eighteen). A significant outcome of the F-111 WPF reinforcement programme was that it resulted in a detailed understanding of the design/assessment tools needed for designing damage tolerant repairs to meet the stringent damage tolerant requirements inherent in both JSSG2006 [1] and FAA ac-120-107b [39]. It is also shown that the in-service performance of the WPF reinforce- ment reveals the shortcomings in the approach outlined in the US Compos- ite Materials Handbook CMH-17-3G [40] for the design of bonded joints. Chapter Eleven presents work performed by the FAA’s Airworthiness Assurance Center at Sandia National Labs in conjunction with Boeing and Federal Express, to validate and introduce composite doubler repair technology to the US commercial aircraft industry. This chapter focuses on repairs to Boeing DC-10 and MD-11 aircraft and builds on the founda- tion of a prior successful Lockheed L-1011 door corner repair [41]. The DC-10 and MD-11 repair programmes highlight the design and analysis, which is based on the design methodology outlined in Chapters Eight and Nine, performance (durability, flaw containment, reliability), installa- tion, and nondestructive inspection issues. This multiyear pilot programme is particularly important in that it was perhaps the first to demonstrate the successful and routine use of composite doubler repairs on commercial aircraft. In this context Chapter Eleven, when taken together with the work presented in Chapters Eight–Ten, further substantiates the extended PABST design methodology [35] whereby the adhesive strain energy density should be kept beneath its fatigue threshold Wf. In 2009, the US FAA introduced a slow growth approach for certifying composite and adhesively bonded structures [39]. This approach requires that delamination and/or disbond growth under cyclic-fatigue loads be slow, stable and predictable. To this end, Chapter Twelve addresses the chal- lenge of developing a methodology capable of enabling this slow growth approach to certification to be implemented. As a result Chapter Twelve examines the growth of disbonds from small naturally occurring material dis- continuities in adhesively bonded structures. It is shown that, for the exam- ples studied, the disbond growth histories can be accurately computed using a form of the Hartman and Schijve variant of the NASGRO crack-growth equation discussed in Chapter Four with ΔK replaced by Δ√G. It is also 7 Rationale and Organisation
- 17. shown that thescatter in the disbond growth histories can be captured by allowing for small changes in the fracture-mechanics threshold term Δ√Gthr. These findings suggest that the Hartman–Schijve crack-growth equation has the potential to address the ‘slow growth’ approach to certify- ing composite/bonded structures and bonded repairs outlined in the US FAA Airworthiness Advisory Circular No: 20-107B [39]. The growth of disbonds in adhesively bonded repairs represents only one of the possible failure mechanisms. Externally bonded composite repairs can also fail by delamination. Indeed, the growth of delaminations in polymer– matrix fibre composites under cyclic-fatigue loading in operational aircraft structures has always been a very important factor. It has the potential to affect significantly the service life and is central to the damage tolerant assess- ment of both composite repairs and composite airframes [1,35,42]. Until recently, the recommended approach was to adopt a ‘no growth’ design [40]. However, the introduction by the FAA of a ‘slow growth’ approach to the certification of composite and bonded structures [39] has focused attention on the experimental data, the analytical tools needed to assess the growth of delaminations under fatigue loads and on means for determin- ing a lower bound on the fatigue threshold. As such the focus of Chapter Thirteen is to address the topic of the growth of delaminations in polymer–matrix fibre composites under cyclic-fatigue loading using a fracture mechanics approach. Specific attention is given to the test and data-reduction procedures that are required in order to determine a ‘valid’ rate of crack growth, da/dN, versus the range of the energy release rate, ΔG (or the maximum strain-energy release rate, Gmax, in a cycle) relationship. This is particularly important since only such ‘valid’ relationships can be reli- ably used (a) to characterise and compare different types of composites and adhesives and (b) for lifting the in-service behaviour of composite structures and bonded composite repairs. To this end, Chapter Thirteen highlights the important role played by fibre bridging that may occur behind the tip of the advancing delamination, and which causes retardation of the fatigue crack growth rate. It is shown that such retardation effects cannot usually be avoided when using the Mode I double-cantilever beam test to ascertain experimentally the fatigue behaviour of composites. Consequently, a means of estimating a ‘valid’ (low retardation) relationship is needed. Chapter Thirteen presents one such approach that is based on the use of the Hartman–Schijve delamination growth equation discussed in Chapters Five and Twelve. This approach also enables a lower bound on the fatigue 8 Rhys Jones et al.
- 18. threshold to bedetermined. This ‘lower bound’ is particularly important if the design is to be based on a no growth approach to delamination or disbonding. Additive manufacture (AM) is increasingly being referred to as the next industrial revolution in that 3D complex shapes can be built layer by layer through material deposition. AM is fast approaching a ‘mainstream’ technol- ogy status [43]. In the aerospace industries, companies that are expanding their 3D manufacturing and production engagement are general electric for jet engines and Lockheed Martin and Boeing for aerospace and defense [43]. Additive metal technologies (AMTs) are part of the additive manufacturing family. In aerospace, AMT can be used for a range of sustainment applications as well as for the creation of replacement parts. In this context, Chapters Fourteen–Eighteen focuses on: SPD, also known as CS, LAD and multiplicative manufacturing. To help clarify the basics of AMT, Chapter Fourteen introduces SPD. SPD, also known as ‘cold spray (CS)’, is a materials consolidation process whereby micron-sized particles of a metal, ceramic and/or polymer are accelerated through a spray gun fitted with a De Laval rocket nozzle using a heated high-pressure gas (i.e., helium or nitrogen) such that the particles exit at supersonic velocities and consolidate on impacting a suitable surface to form a coating or a near-net-shaped part by means of ballistic impinge- ment [44–49]. The particles utilized are typically in the form of commer- cially available powders, ranging in diameter from about 5 to 100 μm, and are accelerated at velocities from 300 to 1500 m/s by injection into a high-velocity stream of gas. Particles that are 5 μm in diameter do not have enough momentum to leave the gas stream and impact the substrate. Hence, nanoparticles cannot be deposited using CS in this manner. Instead agglom- erated nanoparticles or nanostructured powder within the feasible size range are used and CS has been shown to produce nanostructured coatings, as well as nanostructured bulk materials [50]. Chapter Fourteen also briefly discusses the advantages and potential of SPD for both geometry restoration and as a structural repair. Chapter Fifteen builds on this introduction to SPD. It notes that for the new and next generation aircraft, utilisation of computer-aided design tools and technologies and the concurrent computer-aided manufacturing pro- gramming of highly sophisticated machines will often result in optimised designs that significantly reduce and even eliminate any excess material to reduce the overall aircraft weight and thereby enhance operational perfor- mance. Chapter Fifteen further remarks that one outcome of this ‘optimised’ 9 Rationale and Organisation
- 19. design process isthat these new and next generation components will often not be able to be repaired/recovered through the conventional subtractive processes (blending) nor will they be able to be recovered using external mechanically fastened repairs. Hence the challenge is to develop additive metal solutions so that components and/or structure can be returned to operational service in a safe, timely and cost-effective manner. In this context, Chapter Fifteen notes that the maintenance of an aircraft fleet is an ongoing and evolving process and involves both the repair of leg- acy aircraft and the development of capability for next generation aircraft. For legacy aircraft maintenance procedures identify components and struc- ture that no longer meet specification or are considered unfit for purpose and are typically replaced or require extensive repairs. In many instances, the fail- ure to meet specification is due to the loss of metal from the component or structure. This loss can occur through corrosion, wear, erosion or impact. These replaced components or repaired structures can have a significant price tag and adversely affect aircraft availability and as such contribute sig- nificantly to the cost of ownership of the fleet. In the military sphere it has already been noted that the June 2007 Report to Congress by the Under Secretary of the DoD (Acquisition, Technology and Logistics) [25] estimated the cost of corrosion associated with US DoD systems to be between $10 billion and $20 billion annually. This report out- lined the need for research into four primary areas one of which was: Repair processes that restore materials to an acceptable level of structural integrity and functionality. In a recent independent report, the annual cost of corro- sion (2013) for the Australian Defense Aerospace sector was quoted as A$228 million [50]. To meet this challenge Chapter Fifteen focuses on two AMTs, namely SPD and LAD, which are currently being used in the sustainment of aero- space vehicles to reduce the cost of ownership and increase operational availability. Having introduced the concept of SPD (CS) and how it can be can be used to restore geometry and having noted that in the past decades, partic- ularly since the Aloha Airlines incident 1989 [3], Chapter Sixteen notes that corrosion has become one of the primary considerations in both military and commercial aircraft. Indeed, numerous authors, for example, Refs. [51–54], have noted that corrosion can impact heavily the economics, maintenance and safety of aircraft fleets. A number of different types of corrosion have been detected on aircraft structural aluminium alloys with exfoliation, pitting, intergranular corrosion 10 Rhys Jones et al.
- 20. (IGC) and stressvorrosion cracking (SCC) being the prominent types of corrosion generally occurring in structural components exposed to corrosive environments [52,54–56]. This corrosive degradation can result in a signif- icant adverse impact on structural integrity through the loss of material and/ or the generation of cracking. Consequently, before discussing the ability of SPD to restore structural integrity Chapter Sixteen presents a brief discussion on corrosion in aircraft structures. Chapter Sixteen then shows how AMT, specifically SPD (CS), can be used to alleviate the effect of environmental degradation on the structural integrity of wings and horizontal stabilisers. Particular attention is paid to (a) The ability of SPD to essentially restore the load carrying capacity of upper wing skins with skin corrosion and rib stiffened upper wing planks with SCC in the risers. (b) The ability of SPD repair corrosion on the lower (wing skin) tension surfaces. (c) The ability of SPD to alleviate the deleterious effect of IGC at fastener holes on the lower (wing skin) tension surfaces. (d) The ability of the Hartman–Schijve crack growth equation, which is discussed in Chapter Four, to capture the growth of cracks in structures repaired using SPD. Chapter Seventeen builds on the work presented in Chapter Sixteen. To this end, Chapter Seventeen summarises a building block approach that has been used to demonstrate the potential of SPD for repairing and enhancing the airworthiness and integrity of fuselage skins, fuselage lap joints and centre barrel structures. The initial focus is on the application of SPD doublers to fuselage lap joints representative of a Boeing 737 civil transport aircraft and as such complements the work presented in Chapter Nine, which dis- cussed composite repairs to Boeing 737 fuselage lap joints. These successful tests are then supported by an application to an F/A-18 Hornet wing attach- ment centre barrel laboratory test article. It is noted that the centre barrel was subsequently tested for more than three design lifetimes without detectable evidence of degradation (disbonding, delamination or cracking) to the SPD doublers, including those over fastener holes, or the surrounding structure. Chapter Seventeen also shows that even though, during fastener removal to install a Canadian modification mid-way through the test, the SPD was damaged there was no cracking, delamination or disbonding induced and the SPD, and the surrounding structure, remained intact even after cycling for a further 13,000 flight hours, which corresponds to in excess of 2 design lifetimes. This finding highlights the damage tolerant nature of SPD. As such 11 Rationale and Organisation
- 21. Chapter Seventeen showsSPD to be highly durable and damage tolerant and that it can be used to restore structural integrity of full-scale aircraft compo- nents under a representative fighter aircraft spectrum. Chapter Seventeen also shows that SPD doublers can withstand cyclic stress amplitudes of more than 200 MPa without disbonding, delamination or cracking. This is an important finding since it represents a realistic upper bound on the stresses seen on the exterior wetted surface of many operational aircraft. Whereas Chapters Fourteen–Seventeen focused on AMT we have seen that Chapter Ten addressed the use of structural optimisation in conjunction with an externally bonded composite repairto addresssuccessfully theproblem of cracking in primary structure. The success of this combined approach sug- geststhattoaddressproblemsassociatedwithaircraftsustainmentitmaybepos- sible to use multiplicative manufacturing, which is defined as the simultaneous application of subtractive and additive manufacturing processes, to extend the fatigue life of an operational aircraft. To this end, Chapter Eighteen illustrates the potential of multiplicative manufacturing via the LAU-7 missile launcher housing [57]. In this context, Chapter Eighteen show that applying SPD to the external surface of the LAU-7 missile launcher housing further reduces the peak stress at the critical location, in comparison to only adopting a shape optimizationsolution,by29%.Itisshownthatthisadditionalreductioninthe stress field would enable the allowable crack size to be increased from 0.5 to 2.14 mm and would effectively triple the time to the allowable crack size. As such Chapter Eighteen illustrates the potential for using multiplicative manufacturing to extend the fatigue life and maintain the capability of opera- tional aircraft. REFERENCES [1] JSSG-2006, Department of Defense Joint Service Specification Guide, Aircraft Struc- tures, 1998, October. [2] P.C. Miedlar, A.P. Berens, A. Gunderson, J.P. Gallagher, Analysis and support initiative for structural technology (ASIST), 2003. AFRL-VA-WP-TR-2003-3002. [3] C.F. Tiffany, J.P. Gallagher, C.A. Babish IV, Threats to Aircraft Structural Safety, Including a Compendium of Selected Structural Accidents/Incidents, Aeronautical Systems Center, Engineering Directorate, Wright-Patterson Air Force Base, Ohio, 2010. ASC-TR-2010-5002. [4] R. Eastin, ‘WFD’—what is it and what’s ‘LOV’ got to do with it? Int. J. Fatigue 31 (2009) 1012–1016. [5] A.P. Berens, P.W. Hovey, D.A. Skinn, Risk analysis for aging aircraft fleets-volume 1: analysis, WL-TR-91-3066, Flight Dynamics Directorate, Wright Laboratory, Air Force Systems Command, Wright-Patterson Air Force Base, October, 1991. [6] J.W. Lincoln, R.A. Melliere, Economic life determination for a military aircraft, AIAA J. Aircr. 36 (5) (1999) 737–742. 12 Rhys Jones et al.
- 22. [7] J. Schijve,in: C.E. Harris (Ed.), Fatigue life until small cracks in aircraft structures: dura- bility and damage tolerance. Advanced structural integrity methods for airframe dura- bility and damage tolerance, Proceedings FAA/NASA International Symposium, Hampton, 1994, NASA Conference Publication, vol. 3274, 1994, pp. 665–681. [8] R. Jones, Fatigue crack growth and damage tolerance, Fatigue Fract. Eng. Mater. Struct. 37 (2014) 463–483. [9] L. Molent, J.P. Gallagher, The equivalence of EPS and EIFS based on the same crack growth life data, Int. J. Fatigue 80 (2015) 162–170, https://doi.org/10.1016/j. ijfatigue.2015.05.009. [10] W.R. Hendricks, in: S.N. Atluri, S.G. Sampath, P. Tong (Eds.), The Aloha-Airlines Accident—A New Era for Aging Aircraft, Structural Integrity of Aging Airplanes, Springer-Verlag, Berlin, 1991, pp. 153–165. [11] http://www.dailymail.co.uk/news/article-1374574/Southwest-Airlines-fully- operational-plane-cracks-repaired.html. [12] http://www.bbc.co.uk/news/world-us-canada-12954335. [13] FAA-2010-1280-001, Rules and regulations, Fed. Regist. 76 (2011) 6. [14] R.G. Eastin, W. Sippel, in: The “WFD rule”—have we come full circle? Proceedings, 2011 USAF Aircraft Structural Integrity Conference, 29th November–1st December 2011, San Antonio, TX, USA, 2011. [15] L. Molent, S.A. Barter, R.J.H. Wanhill, The lead crack fatigue lifing framework, Int. J. Fatigue 33 (2011) 323–331. [16] L. Molent, R. Jones, A stress versus crack growth rate investigation (aka stress-cubed rule), Int. J. Fatigue 87 (2016) 435–443. [17] J. Ayling, A. Bowler, G. Brick, M. lgnjatovic, Practical application of structural repair fatigue life determination on the AP-3C Orion platform, Adv. Mater. Res. 891– 892 (2014) 1065–1070. [18] B. Main, Structural analysis methodology–F/A-18, Issue 2 AI2, RAAF, Directorate General Technical Airworthiness, December, 2007. [19] J. Duthie, E. Matricciani, DSTO P-3 repair assessment methodology (RAM) coupon testing to validate the cubic rule, DSTO-CR-2010-0367, July, 2011. [20] R. Jones, W. Hu, A.J. Kinloch, A convenient way to represent fatigue crack growth in structural adhesives, Fatigue Fract. Eng. Mater. Struct. 38 (2015) 379–391. [21] R. Jones, A.J. Kinloch, W. Hu, Cyclic-FCG in composite and adhesively-bonded structures: the FAA slow crack growth approach to certification and the problem of similitude, Int. J. Fatigue 88 (2016) 10–18. [22] R. Jones, S. Pitt, A.J. Brunner, D. Hui, Application of the Hartman-Schijve equation to represent mode I and mode II fatigue delamination growth in composites, Compos. Struct. 94 (2012) 1343–1351. [23] A. Mujtaba, S. Stelzer, A.J. Brunner, R. Jones, Influence of cyclic stress intensity thresh- old on the scatter seen in cyclic mode I fatigue delamination growth in DCB tests, Compos. Struct. 169 (2017) 138–143. [24] R.C. Kinzie, D.T. Peeler, in: Structural damage management: new rules, new tools, Proceedings Eighth Joint NASA/FAA/DoD Conference on Aging Aircraft, 31 January 2005, Palms Spring, CA, USA, 2005. [25] Office of the Secretary of Defense, Efforts to reduce corrosion on the military equip- ment and infrastructure of the Department of Defense, Department of Defense Report, June, 2007. [26] V.K. Champagne, M.D. Trexler, B.M. Gabriel, D.J. Helfritch, P.F. Leyman, in: Develop- mentandImplementationcoldspraycoatingsforrepairandcorrosionprotection,Proceed- ings Navy Corrosion Conference, San Diego, 7th–10th, June, 2010. [27] A.A. Baker, R. Jones, Bonded Repair of Aircraft Structures, Martinus Nijhoff, Dordrecht, 1988. 13 Rationale and Organisation
- 23. [28] A.A. Baker,L.R.F. Rose, R. Jones, Advances in the Bonded Composite Repair of Metallic Aircraft Structure, Vols. I and II, Elsevier, Cambridge, MA, ISBN 0-08- 042699-9, 2002. [29] C. Guijt, J. Mazza, Case history: chapter 31 F-16 fuel vent hole repairs, in: A.A. Baker, L.F.J. Rose, R. Jones (Eds.), Advances in the Bonded Composite Repair of Metallic Aircraft Structure, Elsevier Science, Oxford, 2002, pp. 871–884. [30] W.H. Schweinberg, J.W. Fiebig, Case history composite repairs of USAF C-141 and C130 aircraft, in: A.A. Baker, L.F.J. Rose, R. Jones (Eds.), Advances in the Bonded Composite Repair of Metallic Aircraft Structure, Elsevier, Cambridge, MA, 2002, pp. 1009–1032 (Chapter 41). [31] A.A. Baker, R.J. Callinan, M.J. Davis, R. Jones, J.G. Williams, Repair of mirage III aircraft using BFRP crack patching technology, Theor. Appl. Fract. Mech. 2 (1) (1984) 1–16. [32] A.A. Baker, R.J. Chester, M.J. Davis, J.D. Roberts, J.A. Retchford, Reinforcement of the F-111 wing pivot fitting with a boron/epoxy doubler system—materials engineer- ing aspects, Composites 24 (6) (1993) 511–521. [33] L. Molent, R.J. Callinan, R. Jones, Design of an all boron epoxy doubler for the F-111C wing pivot fitting: structural aspects, Compos. Struct. 11 (1) (1989) 57–83. [34] L.J. Hart-Smith, Adhesively bonded double lap joints, NASA CR 112235, January, 1973. [35] D.L. Potter, Primary adhesively bonded structure technology (PABST): design handbook for adhesive bonding, USAF Technical Report, AFFDL-TR-79-3129, November, 1979. [36] L.M. Butkus, J.A. Gaskin, J.M. Greer Jr., C.M. Guijt, N.J. Jacobs, D.F. Kelly, J. J. Mazza, Bonded boron patch repair evaluation: final report, USAFA-TR-2007-06, October, 2007. [37] K.F. Walker, L.R.F. Rose, Case history F-111 lower wing skin repair substantiation (Chapter 27), in: A. Baker, L.R.F. Rose, R. Jones (Eds.), Advances in the Bonded Composite Repair of Metallic Aircraft Structure, Elsevier, Cambridge, MA, ISBN 0-08-042699-9, 2002. [38] Atluri SN., Park JH., Punch EF., O’Donohue PE. and Jones R., Composite repairs of cracked metallic aircraft, Federal Aviation Administration, contract report, DOT/ FAA/CT-92/32, May, 1993. [39] Federal Aviation Authority, Airworthiness Advisory Circular No: 20-107B. Composite Aircraft Structure, 2009. 09/08/2009. [40] CMH-17-3G, Composite Materials Handbook, Volume 3: Polymer Matrix Composites Materials Useage, Design and Analysis, Published by SAE International, Pennsylvania, March 2012. [41] D. Roach, Case history: composite doubler installation on an l-1011 commercial aircraft (Chapter 28), in: A. Baker, L.R.F. Rose, R. Jones (Eds.), Advances in the Bonded Composite Repair of Metallic Aircraft Structure, Elsevier, Cambridge, MA, 2002. 0-08-042699-9. [42] C. Kourloufas, J. Wittkopp, F/A-18A/B inner wing bonded joint defects—through life management strategy, ADF Aircraft Structural Integrity Symposium (AASIS16), 16th– 18th November, 2016. Available from: http://www.defence.gov.au/DASP/ ConferencesAndSeminars/Presentations.asp. [43] R. D’Aveni, The 3-D printing revolution, Harv. Bus. Rev. (May) (2015). [44] A. Papyrin, Cold spray technology, Adv. Mater. Process. (September) (2001) 49. [45] T.H. Van Steenkiste, Kinetic spray coatings, Surf. Coat. Technol. 111 (1999) 62. [46] T. Stoltenhoff, H. Kreve, H. Richter, An analysis of the cold spray process and its coat- ings, J. Therm. Spray Technol. 11 (4) (2002) 542. 14 Rhys Jones et al.
- 24. [47] Moridi A.,Hassani-Gangaraj SM., Guagliano M. and Dao M., Cold spray coating: review of material systems and future perspectives. Surf. Eng., 2014 VOL 36 NO 6, pgs. 369–395. [48] Dillon T, Champagne V, Trexler M, Consolidation of magnesium alloys using cold spray, PowderMet 2012 Conf., June 10–13, Nashville, TN. [49] L. Ajdelsztajn, Cold Spray and GE Technology, http://www.geglobalresearch.com/ blog/cold-spray-ge-technology. GE Global Research, 2013. 07 November. [50] R. Cottam, V. Luzin, Q. Liu, N. Mathews, Y. Wong, M. Brandt, in: Application of supersonic particle deposition for restoring structural integrity of damaged aircraft struc- tures, 15th Australian International Aerospace Congress (AIAC15), Melbourne, 2013. [51] R. Jones, M. Lo, D. Peng, A. Bowler, M. Dorman, M. Janardhana, N.S. Iyyer, A study into the interaction of intergranular cracking and cracking at a fastener hole, Meccanica 50 (2) (2015) 517–532. [52] M. Liao, T. Mills, Exfoliation corrosion and fatigue modelling, in: Corrosion Fatigue and Environmentally Assisted Cracking in Aging Military Vehicles, 2011, pp. 15-1–15-18. RTO-AG-AVT-140, Research and Technology Organisation of NATO, March. [53] J. Lincoln, in: Corrosion and fatigue: safety issue or economic issue, Proceedings USAF Aircraft Structural Integrity Program-ASIP 98, San Antonio, TX, USA, December 1–3, USAF, 1998. [54] L. Molent, Managing airframe fatigue from corrosion pits—a proposal, Eng. Fract. Mech. 137 (March) (2015) 12–25. [55] R.W. Gould, Repairing corrosion damage, in: Corrosion Fatigue and Environmentally Assisted Cracking in Aging Military Vehicles, 2011, pp. 27-1–27-42, RTO-AG-AVT- 140, Research and Technology Organisation of NATO, March. [56] D.W. Hoeppner, Pitting corrosion: morphology and characterization, in: Corrosion Fatigue and Environmentally Assisted Cracking in Aging Military Vehicles, 2011, pp. 15-1–15-16. RTO-AG-AVT-140, Research and Technology Organisation of NATO, March. [57] M. Heller, J. Calero, S.A. Barter, R.J. Wescott, J. Choi, in: Fatigue life extension pro- gram for LAU-7 missile launcher housings using rework shape optimisation, Proceed- ings Fatigue 2014, Melbourne, Australia, 3rd–7th March, 2014. FURTHER READING [1] https://data.epo.org/publication-server/rest/v1.0/publication-dates/20131002/ patents/EP2644510NWA1/document.pdf. 15 Rationale and Organisation
- 25. CHAPTER TWO Fatigue Requirementsfor Aircraft Structures Russell J.H. Wanhill Emmeloord, The Netherlands NOMENCLATURE ANN artificial neural network ASIP aircraft structural integrity program ASTM American Society for Testing and Materials BB building block CA constant amplitude CFRP carbon fibre reinforced plastic DSG design service goal DSTO Defence Science and Technology Organisation DT damage tolerance EIFS equivalent initial flaw sizes EPFM elastic plastic fracture mechanics EPS equivalent precrack size FAA Federal Aviation Administration FCG fatigue crack growth FSFT full-scale fatigue test(ing) GLARE GLAss REinforced aluminium laminates HELIUM HELIcopter Usage Monitoring database HOLSIP HOListic Structural Integrity Process HUMS Health and Usage Monitoring System IDS initial discontinuity state JSF Joint Strike Fighter LBW laser beam welding LCA light combat aircraft LEFM linear elastic fracture mechanics LOV limit of validity MLB MegaLiner Barrel MSD multiple site fatigue damage MY, MZ vertical and horizontal bending loads NDI nondestructive inspection NLR Nationaal Lucht- en Ruimtevaart Laboratorium NRC National Research Council ODAT operational damage assessment tool OEM original equipment manufacturer QF quantitative fractography Aircraft Sustainment and Repair © 2018 Elsevier Ltd. https://doi.org/10.1016/B978-0-08-100540-8.00002-9 All rights reserved. 17
- 26. QZ ground loads RAAFRoyal Australian Air Force R.H. right hand RUL remaining useful life SHM structural health monitoring S–N stress versus number of cycles USAF United States Air Force VA variable amplitude WFD widespread fatigue damage XWB eXtra Wide Body ΔP pressurisation differential ε–N strain versus number of cycles 1 EVOLUTION OF FATIGUE REQUIREMENTS Since the early 1930s the evolution of aircraft structural integrity has been concerned largely with the service behaviour of high-strength metallic materials, particularly aluminium alloys. Taking a (very) broad view, the his- tory of this evolution is as follows [1,2]: 1930–40: Development of all-metal aircraft for public transport. Design and analysis emphasised static strength, with little consideration of fatigue. 1940–56: Increasing awareness, notably for civil aircraft, of the impor- tance of fatigue for airframe safety. Materials with higher static strengths were developed. However, aluminium alloys did not - and still do not - show corresponding increases in fatigue strength. Design of civil aircraft became based on both static and fatigue strengths (safe-life fatigue design) in 1956. 1956-today: Development of Fail-Safe and Damage Tolerant design methods for both civil and military aircraft. These methods recognise that airframe structures must withstand service loads even when damaged or cracked. Safety should be ensured by fatigue and fracture testing and analysis of damaged structures; preservice and in-service inspections; and eventual repairs, replacements or retirement. Service failures have greatly influenced these developments. In particu- lar, there are six case histories that may be regarded as ‘milestones’ in the aerospace industry’s approach to structural integrity [3–7]. By happenstance these case histories are evenly divided between civil and military aircraft, see Figs 1 [2,8] and 2 [2,9,10]. The significances of these accidents are discussed in Sections 1.1 and 1.2. 18 Russell J.H. Wanhill
- 27. 1.1 Civil AircraftMilestone Accidents De Havilland Comets: In January and April 1954 two Comets suffered pres- sure cabin disintegration resulting from fatigue. Investigation and follow-up gave a general awareness of finite fatigue lives, the significance of Fail-Safety, and the usefulness of full-scale fatigue testing (FSFT). Fail-Safe design principles mean that major parts of the structure are (were) designed firstly to achieve a satisfactory fatigue life with no significant cracking. Secondly, the structure is (was) also designed to be inspectable in service and able to sustain significant and easily detectable damage before 1950 COMETS 1954 DAN-AIR 1977 ALOHA 1988 1960 1970 Fatigue (Safe-Life) Either Fail-Safe Inspection impractical? Yes No Damage tolerance Full-scale fatigue test LOV* 1980 1990 2000 2010 Fig. 1 Milestone civil aircraft accidents and evolution of civil aircraft fatigue require- ments [2,8]: *LOV ¼ limit of validity. 1950 B-47’S 1958 F-111 1969 MB326H 1990 1960 1970 Fatigue (Safe-Life) Static strength requirements only USAF ASIP* Damage tolerance USAF Both Quantitative fractography (QF) and ASIP*: DSTO/RAAF 1980 1990 2000 2010 Fig. 2 Milestone military aircraft accidents and the evolution of military aircraft fatigue requirements [2,9] and techniques [10]: *ASIP ¼ aircraft structural integrity program(mes). 19 Fatigue Requirements for Aircraft Structures
- 28. safety is compromised.These latter requirements were met mainly by struc- tural design concepts having multiple load paths and established residual strength (fracture) requirements in the event of complete or obvious partial failure of one structural element. These Fail-Safe design principles were adopted in 1956, with some exceptions, notably including landing gears, which are not amenable to Fail-Safe design. However, as yet there was no requirement to conduct FSFTs, see Fig. 1. Although FSFT has generally or often been done since the Comet accidents, the requirement for FSFT became mandatory only in 1998, see the last paragraph about the ALOHA Boeing 737 accident. DAN-AIR Boeing 707: In May 1977 a Boeing 707 freighter lost the R. H. horizontal stabiliser owing to fatigue in a supposedly Fail-Safe spar. There were several contributing factors: • inadequate inspection for detecting partial failure of the spar upper chord • inadequate residual strength after complete failure of the upper chord • unanticipated high service loads • stabiliser modification not checked by FSFT. This accident prompted (i) reconsidering the problems of ageing aircraft: it became clear that some inspection methods and schedules were inadequate and required supplementary programmes; and (ii) the need for damage tol- erance (DT), which replaced the Fail-Safe design principles in 1978. In civil airworthiness terms, DT means that fatigue cracks will be detected before the safety is compromised, but unlike Fail-Safety there is no premise that cracking will become obvious before it reduces the residual strength below the required safety level [8]. ALOHA Boeing 737: In April 1988, a Boeing 737 suffered explosive decompression with a loss of part of the pressure cabin, subsequently landing safely. The physical manifestation of the accident was multiple site fatigue damage (MSD) along a critical rivet row of the upper skin lap splice. There were several contributing factors: • skin splice design relied on cold-bonding as well as riveting • bonds susceptible to corrosion and disbonding • neglected inspections and repairs. This accident led to recognition that widespread fatigue damage (WFD), of which MSD is one aspect, can cause loss of Fail-Safety. It also prompted more actions to ensure the safety of ageing aircraft, and corrosion control programmes for both civil and military aircraft. Other instances of WFD led eventually to mandatory FSFT from 1998 onwards; and in 2011 the limit of validity (LOV) concept for aircraft above 20 Russell J.H. Wanhill
- 29. 34,000 kg. TheLOV concept requires FSFT to determine the onset of WFD [11], and as such effectively sets a Safe-Life equivalent for this category of aircraft. 1.2 Military Aircraft Milestone Accidents Boeing B-47s: In March and April 1958 four Boeing B-47s suffered fatigue failures resulting in wing losses. After the fourth accident a fleet recovery programme was initiated. This required development of a structural integ- rity programme that included FSFT. This programme was completed in 1959. Also in 1959, the U.S. Air Force (USAF) adopted Safe-Life fatigue design principles and a permanent aircraft structural integrity program (ASIP), see Fig. 2, that included FSFT. This became a formal requirement (mandatory) in 1969. It should be noted that all ASIPs, including those introduced by the Royal Australian Air Force (RAAF) after the MB-326H accident in 1990 (see below), involve much more than fatigue issues [12]. General Dynamics F-111: In December 1969 a General Dynamics F-111A lost the left wing after only 107 airframe flight hours and at less than half the design limit load. Failure occurred by fast fracture after a small amount of fatigue from a large manufacturing flaw in the lower plate of the L.H. wing pivot fitting. The plate was made from a high-strength steel with limited fracture toughness. This failure and FSFT fatigue and fracture problems led to a fracture con- trol programme for the F-111 critical steel components [13]. Furthermore, the F-111 problems and early fatigue cracking and WFD in Lockheed C-5A wing boxes [14] resulted in the USAF abandoning the Safe-Life policy and introducing their DT approach in 1974–75 [15,16]. This preceded the civil aircraft DT requirements by 3–4 years, cf. Figs 1 and 2. There are some dif- ferences, the main one being that the USAF DT approach requires that initial flaws or cracks should be assumed to be present in new structures. However, there is no universal acceptance of the USAF DT require- ments for military aircraft. For example, the U.S. Navy still uses the Safe- Life approach. Aermacchi MB-326H: In November 1990, an Aermacchi MB-326H lost the left wing owing to fatigue failure in the wing spar at only 70% of the Safe-Life derived from FSFT. Investigation showed that fatigue began from a badly drilled bolt hole. 21 Fatigue Requirements for Aircraft Structures
- 30. Further investigation revealed(i) manufacturing flaws in many holes, (ii) fatigue cracks growing from bolt hole manufacturing flaws in spars from other aircraft, and (iii) fatigue cracks growing from normal quality holes and other structural details [17]. A fleet recovery programme based on teardown of several high-life wings, followed by quantitative fractography (QF) of the growth of in- service fatigue cracks, showed that the original FSFT-based Safe-Life was highly unconservative and that wing replacements were the only feasible option. In addition the RAAF’s structural integrity policy changed similarly to the USAF’s after the B-47 accidents in 1958. Since 1990 the RAAF has established comprehensives ASIPs for each fleet. Some have included QF-based fatigue analyses. 2 CONTINUING DEVELOPMENTS This section summarizes some developments in fatigue analyses and airframe materials that have implications for future aircraft structural integ- rity requirements and ASIPs. 2.1 Fatigue Analyses for Conventional Alloys Ongoing fatigue research aims to improve structural analysis capabilities and methods for fatigue life and crack growth predictions. Two of the main developments are (i) the study of short fatigue crack growth and (ii) the possible and actual effects of corrosion on fatigue and the combined action of corrosion and fatigue. Short crack growth: When the USAF DT approach was first introduced, in the early 1970s, there were few actual data for short crack growth. Owing to advances in QF this is no longer the case, particularly since the MB326H accident in 1990. There are now many data for early crack growth in high- performance aircraft [10], including the types and sizes of fatigue-initiating initial discontinuities [18]. Analysis methods have been developed to use these data for life predictions and reassessments, including the use of equiv- alent precrack sizes (EPS) to represent realistic initial discontinuities [10]. Corrosion and fatigue: Since the ALOHA Boeing 737 accident there has been much effort to develop life-prediction models combining corrosion and fatigue. However, most of these models have been based on laboratory 22 Russell J.H. Wanhill
- 31. testing rather thanservice experience, and have assumed that corrosion would not only initiate fatigue but also accelerate it. Evidence from military aircraft indicates that although corrosion can ini- tiate fatigue, the corrosion contribution is predominantly ground-based, while fatigue occurs mainly during flight [19,20]. At present the question whether corrosion and fatigue act in combina- tion or independently requires further investigation per aircraft type and operational conditions. This issue is important since it can have a major impact on service life management [19,21]. 2.2 Airframe Materials Aluminium alloys have predominated in airframes since introduction of the Boeing 247 (1933) and Douglas DC-2 (1934), but composites (mainly car- bon fibre reinforced plastic, CFRP) and titanium alloys provide competition for certain applications. It is also important to note that the choices of mate- rials for different types of aircraft vary greatly: Transport aircraft: Aluminium alloys still account for about 60% of the airframe structural weight, with the notable exceptions of the Boeing 787 Dreamliner and Airbus A350 XWB, which use 50%–53% composites and only 19%–20% aluminium. Tactical aircraft: The emphasis tends to be on composites and lesser per- centages of titanium, aluminium and steels. For example, the British Aero- space/DASA/CASA/Alenia Eurofighter and Lockheed Martin F-35 both use 35%–40% composites, most of which are in the external skin and substructure. There are exceptions, notably the Lockheed Martin and Boeing F-22, which uses only 24% composites and 20% aluminium. Much of the F-22 structure (42%) is titanium, which can tolerate the relatively high service temperatures, especially in the aft fuselage. Helicopters: Owing to the importance of weight savings for vertical lift aircraft, even higher percentages of composites are being used: e.g. the Bell- Boeing V-22 Osprey airframe consists of nearly 80% composites [22] and the NHIndustries NH90 airframe is over 90% composites. The modern trend is to design and build hybrid airframe structures, not least because of the availability of 3rd-generation aluminium–lithium (Al–Li) alloys. Al–Li alloys provide practical structural efficiencies that can rival those of CFRP composites, resulting in similar weight savings with respect to conventional aluminium alloys. The Airbus A380 represents an important example of a recent hybrid airframe construction, see Fig. 3. 23 Fatigue Requirements for Aircraft Structures
- 32. The structural integrityimplications of using the newer materials are summarised here: (1) Replacement of conventional aluminium alloys by the newer (3rd) generation Al–Li alloys will hardly change the fatigue and fracture structural analysis approaches, since these alloys have reached a similar level of reliability in their engineering properties [23]. (2) GLARE (glass reinforced aluminium laminates) is a composite used extensively in the Airbus A380 pressure cabin, see Fig. 3. GLARE has fatigue and fracture properties amenable to conventional analyses for all-metal airframes, although there are some differences. These have led to a GLARE design ‘toolbox’ to enable compliance with current airworthiness regulations for metallic structures [24]. (3) CFRPs have very different properties and design principles. They have high fatigue resistance when free from defects and stress con- centrations, but they are susceptible to impact damage and subse- quent cracking and delamination. The damage growth—which need not be by fatigue—is difficult to predict. This also makes it difficult to validate repairs. Other related problems are inspection reliability and difficulties in analysing complex components to predict the onset of failure. GLARE upper fuselage panels Al-Li 2099-T83, 2196-T8511 and CFRP floor beams CFRP empennage CFRP pressure bulkhead Newer AA 7XXX (upper wing) and 2XXX (lower wing) alloys; Al-Li 2050 –T84 internal structure; CFRP ribs AA 6XXX alloy LBW lower fuselage panels CFRP centre wing box Fig. 3 Advanced materials and processing for major structural areas of the Airbus A380 [23]: CFRP ¼ carbon fibre reinforced plastic; GLARE ¼ glass reinforced alumi- nium laminates; AA 2XXX, 6XXX and 7XXX ¼ conventional aluminium alloys; Al-Li ¼ aluminium–lithium alloys; LBW ¼ laser beam welding. 24 Russell J.H. Wanhill
- 33. The overall resultis that safety is currently ensured by over- designing CFRPs according to the ‘no growth’ DT principle [25,26]. However, this does not mean that there are no further devel- opments. The Federal Aviation Administration (FAA) actively sponsors research in several key areas, including structural substantiation, DT and maintenance practices, materials control and standardisation, and advanced material forms and processes [27]. These activities continually update the safety and certification requirements for composite struc- tures [27,28]. Looking to the future, there is a current lack of FSFT programmes for CFRP components under realistic conditions. These include the effects of temperature and humidity on the properties. 3 FATIGUE LIFING METHODS FOR METALLIC AIRFRAME STRUCTURES The evolution of fatigue requirements discussed in Section 1 and illus- trated by Figs 1 and 2 has meant that several fatigue lifing methods have evolved, see Table 1. Some comments on the methods are made here. Stress-life (S–N): This method has been used extensively in the Safe-Life approach. The method relies on simple cumulative damage rules to obtain ‘safe’ fatigue lives under constant amplitude (CA) or variable amplitude (VA) load histories. The safe lives are then factored down using scatter factors based on engineering judgement and experience. Strain-life (ε–N): This is a common method used to analyse fatigue dam- age in aircraft fleets, and there are several Original Equipment Manufacturer (OEM) and publicly available tools for doing this. USAF DT: This method is used for many USAF aircraft. It has been effective in ensuring structural safety, whereby the assumed initial flaw or crack sizes (see the F-111 discussion in Section 1.2) are large enough for frac- ture mechanics calculations of FCG using models based on well-established macrocrack growth behaviour. This DT approach has been less satisfactory for analyses of structural dura- bility, i.e. achieving an economic life based on FCG. The initial flaw sizes for durability analyses (equivalent initial flaw sizes, EIFS) were set at 0.254 mm for fastener holes (the most critical locations), with smaller sizes allowed by dem- onstration. These initial crack sizes are well within the short crack regime, where until recently there were few relevant data and reliance had to be made on back-extrapolation of long crack growth behaviour using ‘traditional’ analyt- ical modelling. These issues have been addressed by the DSTO approach. 25 Fatigue Requirements for Aircraft Structures
- 34. Table 1 Surveyof Methods for Fatigue Life Assessment of Metallic Airframe Structures [10] • Stress–life (S–N) fatigue limits, Se; unnotched and notched (Kt); constant amplitude (CA) data modifications to Se mean stress effects (R) linear damage rule, also for variable R scatter factors • Strain–life (ε–N) strain-life equation, unnotched data, R ¼ 1 cyclic stress–strain curve analysis rainflow cycle counting (closed hysteresis loops) stress–strain at critical location (notch analysis) mean stress effects (R) via equivalent strain equations, leading to equivalent strain amplitudes damage accumulation rule • USAF Damage Tolerance (DT) specified equivalent initial flaw sizes (EIFS) based on NDI capabilities back-extrapolation of long crack growth data to derive short crack growth LEFM long crack growth models (noninteraction, yield zone, crack opening, strip yield) to derive variable amplitude (VA) crack growth from constant amplitude (CA) data possible use of the crack opening model for short cracks (FASTRAN); differences in long and short crack thresholds need to be included mainly deterministic: stochastic approach becoming accepted • DSTO approach: implemented by the RAAF actual initial discontinuity/flaw sizes and their EPS actual short-to-long crack growth data using QF data compilations to establish empirical relationships describing crack growth behaviour deterministic (‘upper bound’) estimates of lead crack growth scatter factors • Holistic approach: proposed fatigue nucleation mechanisms (also as functions of notch stress concentrations, Kt) fatigue nucleation lives (S–N and/or ε–N assessments) evaluation and selection of marker load strategies for QF of short-to-long crack growth actual short-to-long crack growth using marker loads and QF establishment, validation and choice of appropriate crack growth models and ‘laws’ deterministic (‘upper bound’) and stochastic approaches environmental effects, notably corrosion 26 Russell J.H. Wanhill
- 35. DSTO approach: Asmentioned in Section 2.1, advances in QF since 1990 have provided many data for early crack growth in high-performance aircraft. Analyses have been developed to use these data for FCG life predictions [10]. Holistic approach: This is a broad programme, see Fig. 4, and is intended for all aircraft types. The inclusion of corrosion is especially noteworthy: see also the remarks about corrosion and fatigue in Section 2.1. The project sponsor is the Canadian National Research Council (NRC), with contribu- tions from other experts from North America, Europe, Australia and Asia ( Japan). 4 FATIGUE LIFING ANALYSES FOR METALLIC AIRFRAME STRUCTURES This section is intended to give generic overviews of fatigue lifing ana- lyses. Much more can be found in specialist books on this topic, for example: • Bannantine, J.A., Comer, J.J., Handrock, J.L., 1998, Fundamentals of Metal Fatigue Analysis, Second Edition, Prentice Hall, Englewood Cliffs, New Jersey. • Schijve, J., 2009, Fatigue of Structures and materials, Second Edition, Springer Netherlands, Dordrecht, the Netherlands. Holistic structural integrity process (HOLSIP) development Holistic life (with all intrinsic/extrinsic factors) P1: Nucleation As-manufactured, IDS Crack/corrosion Short cracks LEFM WFD/MSD LEFM/EPFM Repairable ... ... Damage interaction MSD interaction Fract. toughness Residual strength EPFM/LEFM Special NDI NDI detectable /fretting nucleation Noncontinuum mech. Durability Safe life (no env.) Damage tolerance (no env.) P2: Short crack P3: Long crack P4: Instability Fig. 4 Overview of the HOLSIP programme [29]: IDS ¼ initial discontinuity state; EPFM ¼ elastic plastic fracture mechanics; LEFM ¼ linear elastic fracture mechanics; MSD ¼ multiple site fatigue damage; NDI ¼ nondestructive inspection; WFD ¼ widespread fatigue damage. 27 Fatigue Requirements for Aircraft Structures
- 36. Two especially relevantreports in the present context are: • Hu,W.,Tong,Y.C.,Walker,K.F.,Mongru,D.,Amaratunga,R.,Jackson, P.,2006,A review andassessmentofcurrent lifingmethodologiesandtools in Air Vehicles Division, DSTO Research Report DSTO-RR-0321, Defence Science and Technology Organisation, Melbourne, Australia. • Molent, L., Barter, S.A., Wanhill, R.J.H., 2010, The lead crack fatigue lifing framework, DSTO Research Report DSTO-RR-0353, Defence Science and Technology Organisation, Melbourne, Australia. 4.1 Safe-Life Estimations Stress-life approach: This is the traditional basis for Safe-Life fatigue analysis of aircraft structures. A generic overview of the procedure is shown in Fig. 5. Two points require elaboration: (1) Coupon and component S–N curves are modified into the so-called working curves with the aid of several adjustment factors that are usually proprietary. These factors account for various influences on service components that are not included or experienced by coupons and com- ponents during the laboratory tests. (2) The often-used Pålmgren-Miner cumulative damage rule ignores load sequence effects, and the damage accumulation rate is assumed (incor- rectly) to be independent of stress amplitude. The predicted lives can range from 0.3 to at least 3 the actual life [30]. The potential for unconservative life predictions (values 1) is mitigated by using scatter factors based on engineering judgement and experience, as mentioned in Section 3. Structurally significant components Component stress spectrum Cumulative damage fatigue analysis • Pålmgren-Miner rule to calculate life • Retirement time ≤ safe life Retirement / replacement Structural usage fatigue loads Coupon and component fatigue tests • S–N tests S–N curves S–N safe life on calculated life working curves • Scatter factor Fig. 5 Generic procedure for estimating the fatigue Safe-Life using the stress–life approach. 28 Russell J.H. Wanhill
- 37. Strain-Life approach: Thisis commonly used to analyse fatigue damage in aircraft fleets. The damage (loads and accelerations) data are acquired via strain recorders. There are several OEM and publicly available tools for analysing the data. Fig. 6 shows a generic strain–life procedure to estimate safe fatigue lives and remaining lives from fleet data. In practice the test and service data and cumulative damage analyses are integrated into a computer programme that includes an algorithm for cycle counting. This is usually ‘Rainflow’ cou- nting, which is well described in Bannantine et al. [31]. 4.2 Damage Tolerance Estimations DT approach: A generic analysis approach based on the USAF DT method is shown in Fig. 7. As mentioned in Section 3, this method is used for many USAF aircraft and has been effective in ensuring structural safety. There are two points particularly to note: (1) The FCG analyses use linear elastic fracture mechanics (LEFM) FCG models. These models have difficulties in predicting the behaviour of short fatigue cracks, i.e. cracks with dimensions less than about 0.5 mm [32,33]. (2) Fracture toughness data and residual strength analyses and tests are required to complete the DT procedure. Structurally significant components · εa–N tests εeq.a equation Sl–εl · Strain–life equation · Notch analysis rule damage from εa–N equation · Damage accumulation rule to calculate life · Component stress spectra · Component strain spectra · Cycle counting (hysteresis loops) Cumulative damage fatigue analysis · S–ε cyclic stress–strain curve Structural usage fatigue loads · Scatter factor on calculated life · Retirement time ≤safe life Retirement / replacement safe life Fig. 6 Generic procedure for estimating the safe fatigue life and remaining fatigue life using the strain–life approach. 29 Fatigue Requirements for Aircraft Structures
- 38. Usage flight loads •Static • Fatigue Structural review • Identification of critical areas • Component selection Material fatigue crack growth data • Da/dN versus ΔK (R) • Spectrum loading Stress spectra for critical areas and components Component geometries and stress distributions • Stress intensity factors Fatigue crack growth analysis • Crack growth lives • Crack growth curves, a versus N Residual strength analysis and tests • Critical crack lengths Revision loop (component redesign) NDI inspection techniques • Pre-service and in-service detectable crack sizes Preliminary critical crack lengths • Material fracture toughness data Acceptable inspection intervals Component design acceptable with required inspections No Yes Fig. 7 Generic procedure for estimating acceptable DT FCG properties using linear elastic fracture mechanics (LEFM) FCG models. This procedure is based on the USAF DT fatigue lifing method. 30 Russell J.H. Wanhill
- 39. DSTO approach: Thisdiffers importantly from the DT approach with respect to the required input data and the FCG analysis methods, as may be seen from Table 1. The following points bear repeating with the addition of some important details: (1) The DSTO FCG analyses begin with EPS representing actual initial dis- continuity sizes (‘discontinuity’ is neutral and covers a variety of fatigue- nucleating features; whereas ‘flaw’ is negative and can be inappropriate). Use of actual initial discontinuity data was also the original intention of the USAF DT approach for durability analyses, but until QF advances since 1990 there was a lack of FCG data for (very) early crack growth from initiation sites with EPS values less than about 0.25 mm. Besides being important for durability analyses, service experi- ence has shown the early FCG regime to contribute to improved safety analyses for high-performance aircraft [10,34]. Also, the importance of WFD (notably in large transport aircraft, see Section 1.1) is another motivation for QF-based analyses of early FCG and its development [35]. (2) The intransigence of the early FCG regime with respect to LEFM modelling has shown the need to develop empirical relationships describing the early FCG behaviour [10,34]. In turn, this requires QF data compilations of early FCG from service aircraft, FSFT, and component and coupon tests. Hybrid approaches: Notwithstanding the differences between the USAF DT and DSTO approaches, there is much to be said in favour of a rapprochement. The DSTO’s (now DSTG) position is that both LEFM and empirical FCG modelling can be used where justified by data obtained from service, FSFT and component and coupon tests. Also, the EPS concept for early crack growth has recently [36] been shown to be compatible with the USAF EIFS durability concept mentioned in Section 3. 4.3 Operational Damage Estimations Much effort is currently being put into structural health monitoring (SHM). The evolution of SHM has followed two main strategies: (1) Developing rapid NDI technologies to reduce the inspection time on the ground, often obtaining help from on-board sensing systems. (2) Monitoring of loads or strains or damage or a combination of these to determine the remaining useful life (RUL) and/or setting up inspection intervals for safe operation of the aircraft. Some of these systems have evolved into health and usage monitoring systems (HUMS) which use strain sensors extensively. 31 Fatigue Requirements for Aircraft Structures
- 40. HUMS began andcontinues to be the basic system for helicopters, but it is now being used on fixed-wing aircraft, especially tactical aircraft like the Eurofighter and Indian Light Combat Aircraft (LCA). Strain-sensing equip- ment installed on the aircraft provide operational data that are usually used to estimate the RUL via a Safe-Life analysis, as shown generically in Fig. 6. However, it is also possible to use a HUMS system in a DT analysis to calculate a safe crack growth life [37–39]. One such system, developed by the NLR and called ODAT (operational damage assessment tool), is shown in Fig. 8. The architecture reflects the overall procedure for LEFM Operational Damage Assessment Output module Admin module MatDat module K-fact module CraGro module Specturm module Mission module GeoMat module login module h/c type version External data and tools fleet data representative mission profiles trained models NLR ANN tool all mission profiles tail number main assembly h/c reference data meas. missions subset ANN models HELIUM sub assembly crack characteristics component material type user rights module activity log module dbase mgmt module material data K-solutions crack size stress scale factor stress sequences sequence generator crit det flhrs da/dN = f(ΔK,R) K = f(crack size) projected missions available missions mission mix flight parameter sequences Remaining flight hours OK? yes no Tool Fig. 8 The ODAT system architecture developed by the NLR for a helicopter fa- tigue life monitoring programme: HELIUM ¼ helicopter usage monitoring database; ANN ¼ artificial neural network. (Illustration courtesy of Marcel Bos, NLR.) 32 Russell J.H. Wanhill
- 41. modelling of FCGto determine the number of flight hours needed to grow a fatigue crack from its detected size to the size at which fast fracture will occur. The ODAT system requires a reliable FCG model and knowledge of the following parameters: • the local structural geometry and its effect on the crack driving force (i.e. the so-called β factor in stress intensity factor solutions) • the type of material and product form and the associated FCG properties • the expected operational load history. Several types of test, including CA and VA FCG, and monotonic and cyclic stress–strain tests, are required for the LEFM crack growth modelling. Any desired mission history can be assembled by selecting characteristic flights from the HELIUM database, and an artificial neural network (ANN) derives the local stress responses from the relevant flight parameters. The local stress response sequences can then be used as input to the FCG model. 5 TESTING REQUIREMENTS Accurate and reliable estimates of the fatigue lives of metallic airframes present many challenges, particularly for high-performance aircraft and heli- copters. There is always a demand for lighter structures with reduced manufacturing and operating costs. This leads to relatively highly stressed and efficient designs where fatigue cracking can occur at features such as shallow radii at the junction of flanges, webs and stiffeners, as well as at holes and tight radii. Consequently, there are usually many areas that need to be assessed for their fatigue and FCG lives, and many potential locations at which cracking may occur in service. Engineering fatigue design relies in the first instance on baseline coupon tests to assess the many locations identified as susceptible to cracking. The coupons may be loaded by CA or representative VA load histories, and they may try to represent some features of a built-up structure. The results of these coupon tests are averaged to give an indication of the structural life for a pro- duction aircraft. However, there are significant limitations to this approach: (1) Experience has shown that in high performance aircraft, the compo- nents have many features with the potential to crack, and that each of these features is typical of a single type of ‘representative’ coupon. Hence a component’s average indicated life is equivalent to only the shortest average life from tests on several types of coupons. 33 Fatigue Requirements for Aircraft Structures
- 42. (2) Even whenthe most critical feature of a component has been identified and assessed by coupon testing, the coupons are rarely fully representa- tive, notably with respect to the surface treatments and finishes required for production aircraft. This is important because the commencement of fatigue cracking is primarily surface-influenced and therefore greatly dependent on small surface discontinuities inherent to component pro- duction, as well as any surface-connected discontinuities inherent to the material. These limitations are addressed by other means. One way, which is mandatory for all modern aircraft, is to test actual components, and conduct FSFT on part of the structure or even the full airframe, thereby including the effects of com- ponent geometry and production. (N.B.: FSFT became mandatory for mil- itary aircraft in 1969, and civil aircraft in 1998, see Sections 1.1 and 1.2) Another way is to improve coupon testing by making the coupons opti- mally representative of the most fatigue-critical details, e.g. by applying sur- face treatments and finishes used in component production. This may seem obvious, but it is sometimes neglected or overlooked. 5.1 ‘Building Block’ Testing Procedure Fig. 9 shows a schematic ‘Building Block’ (BB) approach for testing mate- rials, components and structures as part of an aircraft certification process. This is Ground flight tests Components, FSFT Subcomponents Elements Coupons P r e - d e s i g n t e s t s P o s t - d e s i g n t e s t s Fig. 9 ‘Building Block’ fatigue test approach for materials, components and structures: after [40]. 34 Russell J.H. Wanhill
- 43. adapted from aschematic for the Lockheed Martin F-35 Joint Strike Fighter (JSF), but is generically valid. The BB approach may be viewed as a pyramid whose base is the initial material evaluation. Each level of the pyramid is the foundation for the next, and the structural complexity and costs increase with each level up to the FSFT(s). The final phase of certification is ground and flight testing of the aircraft. There are several points to be noted about the testing approach and procedure: Coupon testing level: This includes standard and nonstandard tests. Examples of standard tests are stress–life (S–N); strain–life (ε–N); cyclic stress–strain (S–ε); and LEFM FCG. The latter are used to estimate FCG lives according to the USAF DT fatigue lifing method, see Fig. 7. An appro- priate source for standard tests is ASTM International. Nonstandard coupon tests have been developed for the DSTO approach to DT analysis. One type of coupon is a simple flat specimen with an array of laser-induced crack starter slots. These enable QF measurements of FCG from initial discontinuities 0.05 mm in size. However, QF usually requires ‘marker loads’ to be inserted into the fatigue load histories. The types of markers, and the ways in which they are added, can be chosen to ensure negligible influences of the marker loads on the overall FCG [41]. Element to FSFT levels: At all these levels it is advisable to add marker loads to the fatigue load histories. Sometimes a realistic load history will result in natural crack front markers, but it is better to make sure that they occur. There are comprehensive guidelines for this [41]. Crack front marking is especially relevant to QF analyses of fatigue cracks detected during FSFT teardowns, see Fig. 10. Besides being used (and required) directly for certification, the FSFT teardown results are important for possible design modifications, verifying structural analyses, and deter- mining whether retrofits may be advisable or necessary. FSFT types: It is important to note that FSFT does not imply testing of the entire airframe. This may be done in some instances, but the available space and testing equipment often dictate that major parts of the airframe are tested separately. An example of part-structure FSFT is the Airbus A380 MegaLiner Barrel (MLB) fatigue test, in which candidate fuselage skin materials were tested. Fig. 11 shows (i) the MLB; (ii) the types of applied loads, namely fuselage pressurisation (ΔP) and bending (MY, MZ) and ground loads (QZ); and (iii) the number of simulated flights applied during testing. This is slightly more than twice the nominal Design Service Goal (DSG) of 20,000 flights. 35 Fatigue Requirements for Aircraft Structures
- 44. FSFT requirements: Asmentioned in Section 1.1, cases of WFD in civil transport aircraft led eventually to mandatory FSFT from 1998 onwards; and in 2011 the LOV concept for aircraft above 34,000 kg. The LOV concept requires FSFT to determine the onset of WFD [11]. These changes were reflected in the FSFT requirements, or rather expec- tations. From 1998 to 2011 the expectation was that an FSFT would be done Fig. 10 Airframe fatigue certification and the central role of fatigue testing: see Fig. 9 also: original illustration courtesy of Madeleine Burchill and Simon Barter, DSTG, Melbourne. MZ MZ QZ ΔP 45,402 simulated flights with conservatively high fatigue loads MY MY Fig. 11 Airbus A380 MegaLiner Barrel (MLB) full-scale fatigue test ‘specimen’, the gen- eral loading conditions, and the number of simulated flights applied during testing [42]. 36 Russell J.H. Wanhill
- 45. to a minimumof two DSGs followed by specific inspections and analyses [8]. With introduction of the LOV concept it is recommended that the FSFT is run to three DSGs, followed by residual strength testing [8]. The USAF has followed a similar evolutionary path in its DT approach. Thus MIL-STD-1530C [43], which predates the LOV concept, suggests that the FSFT be extended beyond two DSGs because the onset of WFD does not always occur within the two-lifetime test period [6]. 6 SUMMARY The fatigue requirements for aircraft structures have evolved since the 1950s. However, this has not been a gradual process: ‘milestone’ accidents have caused paradigm shifts for both civil and military aircraft, as may be seen from Figs 1 and 2. The requirements encompass fatigue design and analysis, and verification using several levels of testing. The ‘final’ verification is by FSFT and tear- downs, which provide checks on designs and analyses, and can reveal the need for modifications and retrofits. Besides ‘final’ verification of structural integrity, FSFT and teardowns provide references for any fatigue problems that arise during service. They are also references for investigating whether it is feasible to extend the service life beyond the original design life. This is a not infrequent issue. ACKNOWLEDGEMENTS Thanks are due to Simon Barter, Lorrie Molent and Madeleine Burchill, DSTG, Melbourne, Australia; and Marcel Bos, NLR, Emmeloord, the Netherlands. REFERENCES [1] M.C.-Y. Niu, Airframe Structural Design, Conmilit Press, Wanchai, Hong Kong, 1988, (Chapter 1.0, p. 3). [2] P. Safarian, Fatigue and Damage Tolerance Requirements of Civil Aviation, Lesson 01—Introduction, Winter 2014, University of Washington, Seattle, Washington, 2013. [3] J. Schijve, Fatigue of aircraft materials and structures, Int. J. Fatigue 16 (1) (1994) 21–32. [4] A.F.Blom,Fatiguescienceandengineering—achievementsandchallenges,in:J.Rouchon (Ed.),18thPlantemaMemorialLecturein‘ICAF’2001:DesignforDurabilityintheDigital Age, vol. I, C epaduès-Editions, Toulouse, France, 2002, pp. 3–64. [5] A.J. McEvily, Metal Failures: Mechanisms, Analysis, Prevention, Wiley, New York, 2002, (Chapter 1, pp. 6–8, 13–15). [6] C.F. Tiffany, J.P. Gallagher, C.A. Babish IV, Threats to Aircraft Structural Safety, Including a Compendium of Selected Structural Accidents/Incidents, Aerospace Sys- tems Center Technical Report ASC-TR-2010-5002, Wright-Patterson Air Force Base, Dayton, OH, 2010. 37 Fatigue Requirements for Aircraft Structures
- 46. [7] R.J.H. Wanhill,L. Molent, S.A. Barter, Milestone case histories in aircraft structural integrity, in: S. Hashmi (Ed.), Reference Module in Materials Science and Materials Engineering, Elsevier Inc., Oxford, UK, 2015, pp.1–19. https://doi.org/10.1016/ B978-0-12-803581-8.00847-X. [8] R.G. Eastin, W. Sippel, The “WFD rule”—have we come full circle?, USAF Aircraft Structural Integrity Program Conference 2011, November 29–December 1, 2011, San Antonio, TX, 2011. [9] E.S. Wilson, Developments in RAAF aircraft structural integrity management, in: J. M. Grandage, G.S. Jost (Eds.), Estimation, Enhancement and Control of Aircraft Fatigue Performance, vol. II, Engineering Materials Advisory Services, Warley, UK, 1995, pp. 959–970. [10] S.A. Barter, L. Molent, R.J.H. Wanhill, Fatigue life assessment for high performance metallic airframe structures—an innovative practical approach, in: S.-Y. Ho (Ed.), Structural Failure Analysis and Prediction Methods for Aerospace Vehicles and Struc- tures, Bentham E-Books, Bentham Science Publishers, Sharjah, UAR, 2010, pp. 1–17 (Chapter 1). [11] Federal Aviation Administration, Establishing and Implementing Limit of Validity to Prevent Widespread Fatigue Damage, Advisory Circular FAA AC 120-104, U.S. Department of Transportation, Washington, DC, 2011. [12] J.P. Gallagher, A review of the philosophies, processes, methods and approaches that protect in-service aircraft from the scourge of fatigue failures, in: L. Lazzeri, A. Salvetti (Eds.), 21st Plantema Memorial Lecture in ‘Durability and Damage Toler- ance of Aircraft Structures: Metals vs. Composites’, vol. 1, Pacini Ed., Pisa, Italy, 2007, pp. 1–36. [13] W.D. Buntin, Application of fracture mechanics to the F-111 airplane, in: AGARD Conference Proceedings No. 221 on Fracture Mechanics Design Methodology, Advi- sory Group for Aerospace Research and Development, Neuilly-sur-Seine, France, 1977, pp. 3-1–3-12. [14] J.W. Mar, Structural integrity of aging airplanes: a perspective, in: S.N. Atluri, S. G. Sampath, P. Tong (Eds.), Structural Integrity of Aging Airplanes, Springer, Berlin, Germany, 1991, pp. 241–262. [15] MIL-A-83444, Military Specification Airplane Damage Tolerance Requirements, United States Air Force, The Pentagon, VA, 1974. [16] MIL-STD-1530A (11), Military Standard Aircraft Structural Integrity Program, Air- plane Requirements, United States Air Force, The Pentagon, VA, 1975. [17] N. Athiniotis, S.A. Barter, N.T. Goldsmith, Macchi Aircraft Safety-by-Inspection Val- idation Program—A7-076 Starboard Wing Tear-Down Preliminary Report, Defect Assessment Failure Analysis Report No. M21/91, Defence Science and Technology Aeronautical Research Laboratory, Melbourne, Australia, 1991. [18] S.A. Barter, L. Molent, R.J.H. Wanhill, Typical fatigue-initiating discontinuities in metallic aircraft structures, Int. J. Fatigue 41 (1) (2012) 11–22. [19] S.A. Barter, L. Molent, Service fatigue cracking in an aircraft bulkhead exposed to a corrosive environment, Eng. Fail. Anal. 34 (2013) 181–188. [20] L. Molent, Managing airframe fatigue from corrosion pits—a proposal, Eng. Fract. Mech. 137 (2015) 12–25. [21] L. Molent, S.A. Barter, R.J.H. Wanhill, The decoupling of corrosion and fatigue for aircraft service life management, in: A. Siljander (Ed.), ICAF 2015: Structural Integrity: Embracing the Future—Respecting the Past; Supporting Aging Fleets with New Technologies, VTT Technical Research Centre of Finland Ltd, Finland, 2015, pp. 1062–1073. [22] R.B. Deo, J.H. Starnes Jr., R.C. Holzwarth, Low-cost composite materials and struc- tures for aircraft applications, in: Low Cost Composite Structures and Cost Effective 38 Russell J.H. Wanhill
- 47. Application of TitaniumAlloys in Military Platforms, RTO Meeting Proceedings 69 (II), NATO Research and Technology Organisation, Neuilly-sur-Seine, France, 2003, pp. 1-1–1-11. [23] R.J.H. Wanhill, Aerospace applications of aluminum-lithium alloys, in: N.E. Prasad, A. A. Gokhale, R.J.H. Wanhill (Eds.), Aluminum-Lithium Alloys, Processing, Properties and Applications, Butterworth-Heinemann, Elsevier Inc., Oxford, UK, 2013, pp. 503–535. [24] Beumler, T., Flying GLARE® , a contribution to aircraft certification issues in non- damaged and fatigue damaged GLARE® structures, (Doctor’s Thesis), Delft University of Technology, Delft, The Netherlands, 2004. [25] Federal Aviation Administration, Composite Aircraft Structure, Advisory Circular FAA AC-20-107B, Change 1, August 24, 2010, U.S. Department of Transportation, Washington, DC, 2010. [26] MIL-HDBK-17-3F, Military composite materials handbook, in: Polymer Matrix Composites Materials Usage, Design, and Analysis, vol. 3, U.S. Department of Defense, The Pentagon, VA, 2013. [27] Ilcewicz, L., 2007, AC 20-107A “Composite Aircraft Structure” updates, Commercial Aircraft Composite Repair Committee Meeting, Wichita, Kansas, 14 November, 2007. [28] Ilcewicz, L., 2009, Updates to AC 20-107B “Composite Aircraft Structure”, Compos- ite Damage Tolerance Maintenance Workshop, Tokyo, Japan, 5 June 2009. [29] J.P. Komorowski, N.C. Bellinger, M. Liao, A. Fillion, Application of the holistic struc- tural integrity process to Canadian Forces challenges, in: ASIP 2007, USAF Aircraft Structural Integrity Program 2007, December 4–6, 2007, Palm Springs, CA, 2007. [30] W. Sch€ utz, The prediction of fatigue life in the crack initiation and propagation stages— a state of the art survey, Eng. Fract. Mech. 11 (1979) 405–421. [31] J.A. Bannantine, J.J. Comer, J.L. Handrock, Fundamentals of Metal Fatigue Analysis, second ed., Prentice Hall, Englewood Cliffs, NJ, 1998. [32] R.F.W. Anstee, An assessment of the importance of small crack growth to aircraft design, in: Behaviour of Short Cracks in Aircraft Components, AGARD Conference Proceedings No. 328, Advisory Group for Aerospace Research and Development, Neuilly-sur-Seine, France, 1983, pp. 3-1–3-9. [33] R.F.W. Anstee, P.R. Edwards, A review of crack growth threshold and crack propagation rates at short crack lengths, in: Some Considerations on Short Crack Growth Behaviour in Aircraft Structures, AGARD Report No. 696, Advisory Group for Aerospace Research and Development, Neuilly-sur-Seine, France, 1983, pp. 2-1–2-12. [34] L. Molent, S.A. Barter, R.J.H. Wanhill, The Lead Crack Fatigue Lifing Framework, DSTO Research Report DSTO-RR-0353Defence Science and Technology Organi- sation, Melbourne, Australia, 2010. [35] R.J.H. Wanhill, M.F.J. Koolloos, Fatigue and corrosion in aircraft pressure cabin lap splices, Int. J. Fatigue 23 (2001) S337–S347. [36] J. Gallagher, L. Molent, The equivalence of EPS and EIFS based on the same crack growth life data, Int. J. Fatigue 80 (2015) 162–170. [37] J.B. de Jonge, C.W.J. van Lummel, Development of a Crack Severity Index (CSI) for Quantification of Recorded Stress Spectra, NLR Technical Report NLR-TR-86049, National Aerospace Laboratory NLR, Amsterdam, The Netherlands, 1986. [38] H. Lee, S. Park, H. Kim, Estimation of aircraft structural fatigue life using the crack severity index methodology, J. Aircr. 47 (5) (2010) 1672–1678. [39] Aeronautical Systems Center Engineering Directorate, Methodology for Determina- tion of Equivalent Flight Hours and Approaches to Communicate Usage Severity, Structures Bulletin EN-SB-09-001, ASC/EN, Bldg 560, 2530 Loop Road West, Wright-Patterson AFB, OH, 2009. 39 Fatigue Requirements for Aircraft Structures
- 48. [40] D.L. Ball,D.S. Norwood, S.C. TerMaath, Joint Strike Fighter airframe durability and damage tolerance certification, AIAA paper 2006-1867, 47th AIAA/ASME/ASCE/ AHS/ASC Structures, Structural Dynamics, and Materials Conference, 1–4 May 2006, Newport, Rhode Island, 2006. [41] S.A. Barter, R.J.H. Wanhill, Marker Loads for Quantitative Fractography (QF) of Fatigue in Aerospace Alloys, NLR Technical Report NLR-TR-2008-644, National Aerospace Laboratory NLR, Amsterdam, the Netherlands, 2008. [42] R.J.H. Wanhill, D.J. Platenkamp, T. Hattenberg, A.F. Bosch, P.H. de Haan, GLARE teardowns from the MegaLiner Barrel (MLB) fatigue test, in: M.J. Bos (Ed.), ICAF 2009: Bridging the Gap between Theory and Operational Practice, Springer, Netherlands, Dordrecht, the Netherlands, 2009, pp. 145–167. [43] Department of Defense Standard Practice, Aircraft Structural Integrity Program (ASIP), MIL-STD-1530C(USAF), ASC/ENOI, 2530 Loop Road West, Wright- Patterson AFB, OH, 2005. 40 Russell J.H. Wanhill
- 49. CHAPTER THREE Typical Fatigue-Nucleating Discontinuitiesin Metallic Aircraft Structures Simon A. Barter*, Lorrie Molent*, Russell J.H. Wanhill† *Aerospace Division, Defence Science and Technology Group, Australia † Aerospace Vehicles Division, National Aerospace Laboratory NLR, The Netherlands 1 INTRODUCTION Under more or less ideal conditions (highly polished surfaces in homogeneous materials), a significant period of fatigue-induced microstruc- tural damage can precede fatigue crack nucleation. However, production aircraft components and structures have many sources of surface or near- surface discontinuities from which fatigue cracking may very quickly develop in highly stressed locations once loading has commenced, i.e. upon entry of the aircraft into service. The discontinuities include various forms of machining damage (scratches, grooves, burrs, small tears and nicks); etch pits from surface treat- ments (pickling, anodising); porosity, especially in thick aluminium alloy plate and castings; and in the case of aluminium alloys and steels, constituent particles that may themselves be cracked. A fatigue lifing framework using a lead crack concept1 has been devel- oped by the DST Group for metallic primary airframe components [1,2]. Like the original damage tolerance (DT) concept developed by the United States Air Force (USAF), this framework assumes that fatigue cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial crack sizes and deriving early crack growth behaviour from back-extrapolation of growth data for long cracks, the DST framework uses data for real cracks growing from small discontinuities inherent to the materials and component production. To this end, this paper examines the types of discontinuities that act as a nucleus for fatigue cracks in typical metallic airframe structures with a view of informing 1 Lead cracks are fatigue cracks that grow faster (or fastest) in a particular structural area. Aircraft Sustainment and Repair © 2018 Elsevier Ltd. https://doi.org/10.1016/B978-0-08-100540-8.00003-0 All rights reserved. 41
- 50. the reader asto what types of discontinuities should be analysed when full life prediction estimates are made based on crack growth. Table 1 lists examples of fatigue-nucleating discontinuities that are discussed in this paper. These examples are representative of those capable of starting lead fatigue cracks that have been found to limit the fatigue lives of several aircraft types or at least major components of these aircraft. They have been selected from a variety of aircraft types and for various fatigue conditions. Most are for aluminium alloys owing to their widespread use in metallic airframe structures. Most of the discontinuities listed in Table 1 are typical of a normal production environment. The occurrence of poorly finished holes and peening defects can be avoided to some extent by closer production control or changes to the method of application, i.e. laser shock peening instead of conventional peening. However, such improvements are limited owing to the generally small size of the discontinuities, making them hard to detect by Table 1 Overview of Examples of Fatigue-Nucleating Discontinuities Presented Here Discontinuity Sources Specifics Aircraft Type Fatigue Conditions Poorly finished holes Poor drilling Scoring from fastener Poor de-burring Machining tears/nicks Aermacchi MB326 Airbus A380 MLBa Aermacchi MB326 Aermacchi MB326 Aermacchi MB326 Dassault Mirage III Boeing F/A-18A/B Lockheed P3C FSFTb FSFT Service Service Service and FSFT Service Coupon test FSFT Surface treatments Etch pits Etch pits + machining Intergranular ’penetration’ Chemical milling + peening Pickling Peening laps and cuts Boeing F/A-18A/B Boeing F/A-18A/B Boeing B747 G.D.c F-111 Lockheed P3C Boeing F/A-18A/B Coupon tests Service Service Service and FSFT FSFT Coupon tests Porosity Thick plate porosity Boeing F/A-18A/B FSFT Constituent particles Cracked particles Varying shapes Particles + poor machining Boeing F/A-18A/B Dassault Mirage III Coupon test Coupon tests FSFT a MLB ¼ MegaLiner Barrel. b FSFT ¼ full-scale fatigue test. c G.D. ¼ general dynamics (now Lockheed Martin). 42 Simon A. Barter et al.
- 51. preservice NDI. Othertypes of discontinuity are simply unavoidable owing to limitations of the design and production processes, e.g. etch pits from chemical treatments, porosity in castings (which have only limited use in primary airframe structures), and the constituent particles in aluminium alloys and steels. In all these cases, the presence and abundance of the discontinuities, especially combined with high local stresses, result in early fatigue cracking that leads any other fatigue cracking that may occur. The main objectives of this paper are (a) to present examples of the types of discontinuities that actually nucleate fatigue cracks in typical metallic airframe structures and (b) enable designers, production and maintenance organisations to more fully understand the source and nature of what is usually unexpected fatigue cracking. The examples are for cracks from coupon, component and full-scale tests, and also from service aircraft, includ- ing commercial transport aircraft and high performance military aircraft. In every one of the presented examples, the time to fatigue crack nucleatation was found by nondestructive inspection (NDI) or quantitative fractography (QF) to be very short: in fact, so short as to permit the assump- tion of immediate crack growth for lead cracks when applying the lead crack concept [1,2]. 2 EXAMPLES OF DISCONTINUITIES 2.1 Poorly Finished Holes Fastener hole drill or reamer damage, burrs, and poor de-burring practice are common production defects encountered in metallic airframes. Table 2 lists the examples considered here and their sources: Examples 1 and 3–5 are from Aermacchi MB326 trainers, now retired from RAAF service; Example 2 is from a FSFT of the Airbus MegaLiner Barrel (MLB); Example 6 is from a Dassault Mirage III, also withdrawn from RAAF ser- vice; Example 7 is from coupon tests used to support the Boeing F/A-18A/ B lifing programme; Examples 8 and 9 are from a FSFT of an ex-service Lockheed P3C empennage. Example 1. Fig. 1 shows the fatigue cracking, which started on both sides of the fastener hole at the beginning of the FSFT and ultimately failed the wing main spar. On the right-hand side of the hole (as shown in the figure) three cracks (arrowed) grew from gross drill-damage. These cracks interacted and coalesced, and hence grew faster than a single crack growing from a 43 Typical Fatigue-Nucleating Discontinuities in Metallic Aircraft Structures
- 52. drill-induced tear onthe left-hand side. The red arrow on this side indicates the extent of this crack when the right-hand side failed. Quantitative fractography (QF) of the right-hand side cracks showed that they most probably started during the application of the first block of FSFT loading. Table 2 Examples of Fatigue-Nucleating Discontinuities: Poorly Finished Holes Discontinuity Description Component Aircraft Type Source No. Ref. Drilling: gross damage, tearing Drilling: scoring Fastener insertion with scoring Bad de-burring Machining tears AA7075-T6 main spar GLAREa door beam AA7075-T6 main spar AA7075-T6 main spar AISI 4340 carry- through beam AU4SGb spar AA7050-T7451 coupon AA7075-T6 empennage Aermacchi MB326 Airbus MLB Aermacchi MB326 Aermacchi MB326 Aermacchi MB326 Dassault Mirage III Boeing F/A- 18A/B Lockheed P3C FSFT FSFT Service Service Service Service Test FSFT 1 2 3 4 5 6 7 8 [3] [4] [3] [3] [5] [6] [7] [8] a GLARE ¼ glass reinforced aluminium laminates. b AU4SG AA2014-T6. 1 mm Fig. 1 Example of a poorly drilled hole in a main wing spar made from an AA7075-T6 aluminium alloy extrusion. The cracking on the right side of the hole has nucleated from gross drill damage, while the crack on the left side of the hole nucleated from a drill-induced surface tear during machining. The arrow on the far left indicates the position of the fatigue crack on this side when the right ligament failed. 44 Simon A. Barter et al.
- 53. Example 2. Fig.2 shows two linked-up FSFT fatigue cracks in the bore of a GLARE fastener hole in the MLB passenger door beam. Most cracks grew from the 2024-T3 aluminium alloy sheet corners [4], but in this example a crack also started from scoring owing to poor drilling. This result and other evidence of poor drilling led to a production improvement programme for GLARE components in the Airbus A380. Again, posttest QF indicated that the cracks most probably started very early in the FSFT. Example 3. Fig. 3 shows service fatigue cracks starting from scoring along the bore of a fastener hole in a wing main spar. The scoring was due to inser- tion of a blind fastener. The cracks grew more or less evenly on both sides of the hole, and would probably have failed the spar if it had remained in service. Discovery of these cracks led to an extensive teardown programme, and many similarly damaged holes were found [3]. An analysis of fatigue metre data and a correlation of these data with the progression markings on the fatigue fracture surfaces indicated that the cracking had started very early in the life of the aircraft. Example 4. Fig. 4 shows service fatigue cracks growing from a fastener hole in a wing main spar withdrawn from service. The hole had been badly de-burred, resulting in a split burr that started the upper crack and several nicks (arrowed), one of which started the lower crack. These cracks were found during the teardown programme mentioned for Example 3 [3]. 200µm Fig. 2 Example of fatigue cracking nucleation from scoring in the bore of a fastener hole in the MLB GLARE door beam [4]. The cracking is in one of the AA2024-T3 sheets of the GLARE laminate. 45 Typical Fatigue-Nucleating Discontinuities in Metallic Aircraft Structures
- 54. Example 5. Fig.5 shows a magnetic rubber cast of service fatigue cracks growing from machining tears in a fastener hole. This hole was from a steel centre-section lower beam of the wing carry-through frames, and the cracking was detected during a safety-by-inspection programme of all the aircraft in the fleet. This programme indicated that the cracks 1mm Fig. 3 Example of a hole that was damaged by fastener installation in a main wing spar made from an AA7075-T6 extrusion. Service fatigue cracks have nucleated from bore surface scores produced during insertion of a blind fastener. The middle arrow (bore of the hole) indicates one of the scores along the hole bore. 5mm Fatigue crack Fatigue crack Fig. 4 Example of a hole that was poorly finished in a main wing spar made from an AA7075-T6 extrusion. The upper fatigue crack nucleated from a split in a burr at the edge of the hole. The lower fatigue crack nucleated from a nick on the edge of the hole: other nicks are indicated by the arrows. 46 Simon A. Barter et al.
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- 56. Once upon atime, when Brahma-datta was reigning in Benāres, the Bodisat became his private chaplain. At that time certain fishermen were casting their nets into the river. Now a big fish came swimming along playing lustily with his wife. She still in front of him smelt the smell of a net, and made a circuit, and escaped it. But the greedy amorous fish went right into the mouth of the net. When the fishermen felt his coming in they pulled up the net, seized the fish, and threw it alive on the sand, and began to prepare a fire and a spit, intending to cook and eat it. Then the fish lamented, saying to himself; “The heat of the fire would not hurt me, nor the torture of the spit, nor any other pain of that sort; but that my wife should sorrow over me, thinking I must have deserted her for another, that is indeed a dire affliction!” And he uttered this stanza— “’Tis not the heat, ‘tis not the cold, ’Tis not the torture of the net; But that my wife should think of me, ’He’s gone now to another for delight.’” Now just then the chaplain came down, attended by his slaves, to bathe at the ford. And he understood the language of all animals. So on hearing the fish’s lament, he thought to himself: “This fish is lamenting the lament of sin. Should he die in this unhealthy state of mind, he will assuredly be reborn in hell. I will save him.” And he went to the fishermen, and said— “My good men! don’t you furnish a fish for us every day for our curry?”
- 57. “What is thisyou are saying, sir?” answered the fishermen. “Take away any fish you like!” “We want no other: only give us this one.” “Take it, then, sir.” The Bodisat took it up in his hands, seated himself at the river-side, and said to it, “My good fish! Had I not caught sight of you this day, you would have lost your life. Now henceforth sin no more!” And so exhorting it, he threw it into the water, and returned to the city. When the Teacher had finished this discourse, he proclaimed the Truths. At the end of the Truths the depressed monk was established in the fruit of conversion. Then the Teacher made the connexion, and summed up the Jātaka: “She who at that time was the female fish was the former wife, the fish was the depressed monk, but the chaplain was I myself.” END OF THE STORY OF THE FISH AND HIS WIFE.328 No. 35. VAṬṬAKA JĀTAKA. The Holy Quail. “Wings I have that will not fly.”—This the Master told when journeying through Magadha about the going out of a Jungle Fire.
- 58. For once, whenthe Master was journeying through Magadha, he begged his food in a certain village in that land; and after he had returned from his rounds and had finished his meal, he started forth again, attended by the disciples. Just then a great fire arose in the jungle. Many of the monks were in front, many of them behind. And the fire came spreading on towards them, one mass of smoke and flame. Some of the monks being unconverted were terrified with the fear of death; and called out— “Let’s make a counter-fire, so that the conflagration shall not spread beyond the space burnt out by that.” And taking out their fire-sticks they began to get a light. But the others said, “Brethren, what is this you are doing? ‘Tis like failing to see the moon when it has reached the topmost sky, or the sun as it rises with its thousand rays from the eastern quarter of the world; ‘tis like people standing on the beachy shore and perceiving not the ocean, or standing close to Sineru and seeing not that mighty mountain, for you—when journeying along in company with the greatest Being in earth or heaven—to call out, ‘Let us make a counter fire,’ and to take no notice of the supreme, the Buddha! You know not the power of the Buddhas! Come, let us go to the Master!” And they all crowded together from in front, and from behind, and went up in a body near to the Mighty by Wisdom. There the Master stopped, surrounded by the whole body of disciples. The jungle fire came on roaring as if to overwhelm them. It came right up to the place where the Great Mortal stood, and then—as it came within about sixteen rods of that spot—it went out, like a torch thrust down into water, leaving a space of about thirty-two rods in breadth over which it could not pass! Then the monks began to magnify the Teacher, saying;
- 59. “Oh! how marvellousare the qualities of the Buddhas! The very fire, unconscious though it be, cannot pass over the place where the Buddhas stand. Oh! how great is the might of the Buddhas!” On hearing this the Teacher said— “It is not, monks, through any power I have now that the fire goes out on reaching this plot of ground. It is through the power of a former act of mine. And in all this spot no fire will burn through the whole kalpa, for that was a miracle enduring through a kalpa.”329 Then the venerable Ānanda folded a robe in four, and spread it as a seat for the Teacher. The Teacher seated himself; and when he had settled himself cross-legged, the body of disciples seated themselves reverently round him, and requested him, saying— “What has now occurred, O Lord, is known to us. The past is hidden from us. Make it known to us.” And the Teacher told the tale. Long ago the Bodisat entered upon a new existence as a quail in this very spot, in the land of Magadha; and after having been born in the egg, and having got out of the shell, he became a young quail, in size like a big partridge.330 And his parents made him lie still in the nest, and fed him with food they brought in their beaks. And he had no power either to stretch out his wings and fly through the air, nor to put out his legs and walk on the earth. Now that place was consumed year after year by a jungle fire. And just at that time the jungle fire came on with a mighty roar and seized upon it. The flocks of birds rose up, each from his nest, and flew away shrieking. And the Bodisat’s parents too, terrified with the fear of death, forsook the Bodisat, and fled. When the Bodisat, lying there as he was, stretched forth his neck, and saw the conflagration spreading towards him, he thought: “If I had the power of stretching my wings and flying in the air, or of
- 60. putting out mylegs, and walking on the ground, I could get away to some other place. But I can’t! And my parents too, terrified with the fear of death, have left me all alone, and flown away to save themselves. No other help can I expect from others, and in myself I find no help. What in the world shall I do now!” But then it occurred to him, “In this world there is such a thing as the efficacy of virtue; there is such a thing as the efficacy of truth. There are men known as omniscient Buddhas, who become Buddhas when seated under the Bo-tree through having fulfilled the Great Virtues in the long ages of the past; who have gained salvation by the wisdom arising from good deeds and earnest thought, and have gained too the power of showing to others the knowledge of that salvation; who are full of truth, and compassion, and mercy, and longsuffering; and whose hearts reach out in equal love to all beings that have life. To me, too, the Truth is one, there seems to be but one eternal and true Faith. It behoves me, therefore—meditating on the Buddhas of the past and on the attributes that they have gained, and relying on the one true faith there is in me—to perform an Act of Truth; and thus to drive back the fire, and procure safety both for myself, and for the other birds.” Therefore it is said (in the Scriptures)—
- 61. “There’s power invirtue in the world— In truth, and purity, and love! In that truth’s name I’ll now perform A mystic Act of Truth sublime. Then thinking on the power of the Faith, And on the Conquerors in ages past, Relying on the power of the Truth, I then performed the Miracle!” Then the Bodisat called to mind the attributes of the Buddhas who had long since passed away; and, making a solemn asseveration of the true faith existing in himself, he performed the Act of Truth, uttering the verse— “Wings I have that will not fly, Feet I have that will not walk; My parents, too, are fled away! O All-embracing Fire—go back!”331 Then before him and his Act of Truth the Element went back a space of sixteen rods; but in receding it did not return to consume the forest; it went out immediately it came to the spot, like a torch plunged into water. Therefore it is said— “For me and for my Act of Truth The great and burning fire went out, Leaving a space of sixteen rods, As fire, with water mixed, goes out.” And as that spot has escaped being overwhelmed by fire through all this kalpa, this is said to be ‘a kalpa-enduring miracle.’ The Bodisat having thus performed the Act of Truth, passed away, at the end of his life, according to his deeds.
- 62. When the Teacherhad finished this discourse, in illustration of what he had said (“That this wood is not passed over by the fire is not a result, O monks, of my present power; but of the power of the Act of Truth I exercised as a new-born quail”), he proclaimed the Truths. At the conclusion of the Truths some were Converted, some reached the Second Path, some the Third, some the Fourth. And the Teacher made the connexion, and summed up the Jātaka, “My parents at that time were my present parents, but the King of the Quails was I myself.” END OF THE STORY OF THE HOLY QUAIL.332 No. 36. SAKUṆA JĀTAKA. The Wise Bird and the Fools. “The earth-born tree.”—This the Master told when at Jetavana, about a monk whose hut was burned. A certain monk, says the tradition, received from the Teacher a subject for meditation, and leaving Jetavana, took up his abode in a dwelling in a forest near a border village, belonging to the people of Kosala. Now in the very first month his hut was burned down; and he told the people, saying, “My hut is burnt down, and I live in discomfort.” “Our fields are all dried up now,” said they; “we must first irrigate the lands.” When they were well muddy, “We must sow the seed,” said they. When the seed was sown, “We must put up the fences,” was the excuse. When the fences were up, they declared, “There will be cutting, and reaping, and treading-out to do.” And thus,
- 63. telling first ofone thing to be done and then of another, they let three months slip by. The monk passed the three months in discomfort in the open air, and concluded his meditation, but could not bring the rest of his religious exercise to completion. So when Lent was over he returned to the Teacher, and saluting him, took his seat respectfully on one side. The Teacher bade him welcome, and then asked him, “Well, brother, have you spent Lent in comfort? Have you brought your meditation to its conclusion?” He told him what had happened, and said, “As I had no suitable lodging, I did not fully complete the meditation.” “Formerly, O monk,” said the Teacher, “even animals were aware what was suitable for them, and what was not. Why did not you know it?” And he told a tale. Long ago, when Brahma-datta was reigning in Benāres, the Bodisat came to life again as a bird, and lived a forest life, attended by a flock of birds, near a lofty tree, with branches forking out on every side. Now one day dust began to fall as the branches of the tree rubbed one against another. Then smoke began to rise. The Bodisat thought, on seeing this,— “If these two branches go on rubbing like that they will send out sparks of fire, and the fire will fall down and seize on the withered leaves; and the tree itself will soon after be consumed. We can’t stop here; we ought to get away at once to some other place.” And he addressed the flock in this verse:
- 64. “The earth-born tree,on which We children of the air depend, It, even it, is now emitting fire. Seek then the skies, ye birds! Behold! our very home and refuge Itself has brought forth danger!” Then such of the birds as were wise, and hearkened to the voice of the Bodisat, flew up at once with him into the air, and went elsewhere. But such as were foolish said one to another, “Just so! Just so! He’s always seeing crocodiles in a drop of water!” And paying no attention to what he said, they stopped there. And not long afterwards fire was produced precisely in the way the Bodisat had foreseen, and the tree caught fire. And smoke and flames rising aloft, the birds were blinded by the smoke; they could not get away, and one after another they fell into the fire, and were burnt to death! When the Teacher had finished this discourse with the words, “Thus formerly, O monk, even the birds dwelling on the tree-tops knew which place would suit them and which would not. How is it that you knew it not?” he proclaimed the Truths. At the conclusion of the Truths the monk was established in Conversion. And the Teacher made the connexion, and summed up the Jātaka, “The birds who at that time listened to the voice of the Bodisat were the followers of the Buddha, but the Wise Bird was I myself.” END OF THE STORY OF THE WISE BIRD AND THE FOOLS.
- 65. No. 37. TITTIRA JĀTAKA. ThePartridge, Monkey, and Elephant. “’Tis those who reverence the aged.”—This the Master told on the road to Sāvatthi about Sāriputta being kept out of a night’s lodging. For when Anātha Piṇḍika had finished his monastery, and sent word to the Teacher, the latter left Rājagaha and arrived at Vesali; and after resting there a short time, he set out again on the road to Sāvatthi.333 On that occasion the pupils of the Six went on in front, and before lodgings had been taken for the Elders, occupied all the places to be had, saying,— “This is for our superior, this for our instructor, and these for us.” The Elders who came up afterwards found no place to sleep in. Even Sāriputta’s pupils sought in vain for a lodging-place for the Elder. So the Elder having no lodging passed the night either walking up and down, or sitting at the foot of a tree, not far from the place where the Teacher was lodged. In the early morning the Teacher came out and coughed. The Elder coughed too. “Who’s there?” said the Teacher. “’Tis I, Lord; Sāriputta,” was the reply. “What are you doing here, so early, Sāriputta?” asked he. Then he told him what had happened; and on hearing what the Elder said, the Teacher thought,—
- 66. “If the monkseven now, while I am yet living, show so little respect and courtesy to one another, what will they do when I am dead?” And he was filled with anxiety for the welfare of the Truth. As soon as it was light he called all the priests together, and asked them— “Is it true, priests, as I have been told, that the Six went on in front, and occupied all the lodging-places to the exclusion of the Elders?” “It is true, O Blessed One!” said they. Then he reproved the Six, and addressing the monks, taught them a lesson, saying,— “Who is it, then, O monks, who deserves the best seat, and the best water, and the best rice?” Some said, “A nobleman who has become a monk.” Some said, “A Brāhman, or the head of a family who has become a monk.” Others said, “The man versed in the Rules of the Order; an Expounder of the Law; one who has attained to the First Jhāna, or the Second, or the Third, or the Fourth.” Others again said, “The Converted man; or one in the Second or the Third Stage of the Path to Nirvāna; or an Arahat; or one who knows the Three Truths; or one who has the Sixfold Wisdom.”334 When the monks had thus declared whom they each thought worthy of the best seat, and so on, the Teacher said: “In my religion, O monks, it is not the being ordained from a noble, or a priestly, or a wealthy family; it is not being versed in the Rules of the Order, or in the general or the metaphysical books of the Scriptures; it is not the attainment of the Jhānas, or progress in the Path of Nirvāna, that is the standard by which the right to the best seat, and so on, is to be judged. But in my religion, O monks, reverence, and service, and respect, and civility, are to be paid according to age; and for the aged the best seat, and the best water,
- 67. and the bestrice are to be reserved. This is the right standard; and therefore the senior monk is entitled to these things. And now, monks, Sāriputta is my chief disciple; he is a second founder of the Kingdom of Righteousness, and deserves to receive a lodging immediately after myself. He has had to pass the night without a lodging at the foot of a tree. If you have even now so little respect and courtesy, what will you not do as time goes on?” And for their further instruction he said: “Formerly, O monks, even animals used to say, ‘It would not be proper for us to be disrespectful and wanting in courtesy to one another, and not to live on proper terms with one another. We should find out who is eldest, and pay him honour.’ So they carefully investigated the matter, and having discovered the senior among them, they paid him honour; and so when they passed away, they entered the abode of the gods.” And he told a tale. Long ago there were three friends living near a great Banyan-tree, on the slope of the Himālaya range of mountains—a Partridge, a Monkey, and an Elephant. And they were wanting in respect and courtesy for one another, and did not live together on befitting terms. But it occurred to them, “It is not right for us to live in this manner. What if we were to cultivate respect towards whichever of us is the eldest?” “But which is the eldest?” was then the question; until one day they thought, “This will be a good way for finding it out;” and the Monkey and the Partridge asked the Elephant, as they were all sitting together at the foot of the Banyan-tree—
- 68. “Elephant dear! Howbig was this Banyan Tree at the time you first knew it?” “Friends!” said he, “When I was little I used to walk over this Banyan, then a mere bush, keeping it between my thighs; and when I stood with it between my legs, its highest branches touched my navel. So I have known it since it was a shrub.” Then they both asked the Monkey in the same way. And he said, “Friends! when I was quite a little monkey I used to sit on the ground and eat the topmost shoots of this Banyan, then quite young, by merely stretching out my neck. So that I have known it from its earliest infancy.” Then again the two others asked the Partridge as before. And he said— “Friends! There was formerly a lofty Banyan-tree in such and such a place, whose fruit I ate and voided the seeds here. From that this tree grew up: so that I have known it even from before the time when it was born, and am older than either of you!” Thereupon the Elephant and the Monkey said to the clever Partridge — “You, friend, are the oldest of us all. Henceforth we will do all manner of service for you, and pay you reverence, and make salutations before you, and treat you with every respect and courtesy, and abide by your counsels. Do you in future give us whatever counsel and instruction we require.” Thenceforth the Partridge gave them counsel, and kept them up to their duty, and himself observed his own. So they three kept the Five Commandments; and since they were courteous and respectful to one another, and lived on befitting terms one with another, they became destined for heaven when their lives should end.
- 69. “The holy lifeof these three became known as ‘The Holiness of the Partridge.’ For they, O monks, lived in courtesy and respect towards one another. How then can you, who have taken the vows in so well- taught a religion, live without courtesy and respect towards one another? Henceforth, O monks, I enjoin upon you reverence, and service, and respect, according to age; the giving of the best seats, the best water, and the best food according to age; and that the senior shall never be kept out of a night’s lodging by a junior. Whoever so keeps out his senior shall be guilty of an offence.” It was when the Teacher had thus concluded his discourse that he, as Buddha, uttered the verse— “’Tis those who reverence the old That are the men versed in the Faith. Worthy of praise while in this life, And happy in the life to come.” When the Teacher had thus spoken on the virtue of paying reverence to the old, he established the connexion, and summed up the Jātaka, by saying, “The elephant of that time was Moggallāna, the monkey Sāriputta, but the partridge was I myself.” END OF THE STORY OF THE PARTRIDGE, THE MONKEY, AND THE ELEPHANT.335 No. 38. BAKA JĀTAKA. The Cruel Crane Outwitted. “The villain though exceeding clever.”—This the Master told when at Jetavana about a monk who was a tailor.
- 70. There was amonk, says the tradition, living at Jetavana, who was exceeding skilful at all kinds of things that can be done to a robe, whether cutting out, or piecing together, or valuing, or sewing it. Through this cleverness of his he was always engaged in making robes, until he became known as ‘The robe-maker.’ Now what used he to do but exercise his handicraft on some old pieces of cloth, so as to make out of them a robe soft and pleasant to the touch; and when he had dyed it, he would steep it in mealy water, and rub it with a chankshell so as to make it bright and attractive, and then lay it carefully by. And monks who did not understand robe work, would come to him with new cloths, and say — “We don’t understand how to make robes. Be so kind as to make this into a robe for us.” Then he would say, “It takes a long time, Brother, before a robe can be made. But I have a robe ready made. You had better leave these cloths here and take that away with you.” And he would take it out and show it to them. And they, seeing of how fine a colour it was, and not noticing any difference, would give their new cloths to the tailor-monk, and take the robe away with them, thinking it would last. But when it grew a little dirty, and they washed it in warm water, it would appear as it really was, and the worn-out places would show themselves here and there upon it. Then, too late, they would repent. And that monk became notorious, as one who passed off old rags upon anybody who came to him. Now there was another robe-maker in a country village who used to cheat everybody just like the man at Jetavana. And some monks who knew him very well told him about the other, and said to him— “Sir! there is a monk at Jetavana who, they say, cheats all the world in such and such a manner.”
- 71. “Ah!” thought he,“’twould be a capital thing if I could outwit that city fellow!” And he made a fine robe out of old clothes, dyed it a beautiful red, put it on, and went to Jetavana. As soon as the other saw it, he began to covet it, and asked him— “Is this robe one of your own making, sir?” “Certainly, Brother,” was the reply. “Sir! let me have the robe. You can take another for it,” said he. “But, Brother, we village monks are but badly provided. If I give you this, what shall I have to put on?” “I have some new cloths, sir, by me. Do you take those and make a robe for yourself.” “Well, Brother! this is my own handiwork; but if you talk like that, what can I do? You may have it,” said the other; and giving him the robe made of old rags, he took away the new cloths in triumph. And the man of Jetavana put on the robe; but when a few days after he discovered, on washing it, that it was made of rags, he was covered with confusion. And it became noised abroad in the order, “That Jetavana robe-maker has been outwitted, they say, by a man from the country!” And one day the monks sat talking about this in the Lecture Hall, when the Teacher came up and asked them what they were talking about, and they told him the whole matter. Then the Teacher said, “Not now only has the Jetavana robe-maker taken other people in in this way, in a former birth he did the same. And not now only has he been outwitted by the countryman, in a former birth he was outwitted too.” And he told a tale.
- 72. Long ago theBodisat was born to a forest life as the Genius of a tree standing near a certain lotus pond. Now at that time the water used to run short at the dry season in a certain pond, not over large, in which there were a good many fish. And a crane thought, on seeing the fish— “I must outwit these fish somehow or other and make a prey of them.” And he went and sat down at the edge of the water, thinking how he should do it. When the fish saw him, they asked him, “What are you sitting there for, lost in thought?” “I am sitting thinking about you,” said he. “Oh, sir! what are you thinking about us?” said they. “Why,” he replied; “there is very little water in this pond, and but little for you to eat; and the heat is so great! So I was thinking, ‘What in the world will these fish do now?’” “Yes, indeed, sir! what are we to do?” said they. “If you will only do as I bid you, I will take you in my beak to a fine large pond, covered with all the kinds of lotuses, and put you into it,” answered the crane. “That a crane should take thought for the fishes is a thing unheard of, Sir, since the world began. It’s eating us, one after the other, that you’re aiming at!” “Not I! So long as you trust me, I won’t eat you. But if you don’t believe me that there is such a pond, send one of you with me to go and see it.” Then they trusted him, and handed over to him one of their number —a big fellow, blind of one eye, whom they thought sharp enough in
- 73. any emergency, afloator ashore. Him the crane took with him, let him go in the pond, showed him the whole of it, brought him back, and let him go again close to the other fish. And he told them all the glories of the pond. And when they heard what he said, they exclaimed, “All right, Sir! You may take us with you.” Then the crane took the old purblind fish first to the bank of the other pond, and alighted in a Varaṇa-tree growing on the bank there. But he threw it into a fork of the tree, struck it with his beak, and killed it; and then ate its flesh, and threw its bones away at the foot of the tree. Then he went back and called out— “I’ve thrown that fish in; let another come!” And in that manner he took all the fish, one by one, and ate them, till he came back and found no more! But there was still a crab left behind there; and the crane thought he would eat him too, and called out— “I say, good crab, I’ve taken all the fish away, and put them into a fine large pond. Come along. I’ll take you too!” “But how will you take hold of me to carry me along?” “I’ll bite hold of you with my beak.” “You’ll let me fall if you carry me like that. I won’t go with you!” “Don’t be afraid! I’ll hold you quite tight all the way.” Then said the crab to himself, “If this fellow once got hold of fish, he would never let them go in a pond! Now if he should really put me into the pond, it would be capital; but if he doesn’t—then I’ll cut his throat, and kill him!” So he said to him—
- 74. “Look here, friend,you won’t be able to hold me tight enough; but we crabs have a famous grip. If you let me catch hold of you round the neck with my claws, I shall be glad to go with you.” And the other did not see that he was trying to outwit him, and agreed. So the crab caught hold of his neck with his claws as securely as with a pair of blacksmith’s pincers, and called out, “Off with you, now!” And the crane took him and showed him the pond, and then turned off towards the Varaṇa-tree. “Uncle!” cried the crab, “the pond lies that way, but you are taking me this way!” “Oh, that’s it, is it!” answered the crane. “Your dear little uncle, your very sweet nephew, you call me! You mean me to understand, I suppose, that I am your slave, who has to lift you up and carry you about with him! Now cast your eye upon the heap of fish-bones lying at the root of yonder Varaṇa-tree. Just as I have eaten those fish, every one of them, just so I will devour you as well!” “Ah! those fishes got eaten through their own stupidity,” answered the crab; “but I’m not going to let you eat me. On the contrary, it is you that I am going to destroy. For you in your folly have not seen that I was outwitting you. If we die, we die both together; for I will cut off this head of yours, and cast it to the ground!” And so saying, he gave the crane’s neck a grip with his claws, as with a vice. Then gasping, and with tears trickling from his eyes, and trembling with the fear of death, the crane beseeched him, saying, “O my Lord! Indeed I did not intend to eat you. Grant me my life!” “Well, well! step down into the pond, and put me in there.” And he turned round and stepped down into the pond, and placed the crab on the mud at its edge. But the crab cut through its neck as clean as one would cut a lotus-stalk with a hunting-knife, and then only entered the water!
- 75. “When the Geniuswho lived in the Varaṇa-tree saw this strange affair, he made the wood resound with his plaudits, uttering in a pleasant voice the verse— “The villain, though exceeding clever, Shall prosper not by his villany. He may win indeed, sharp-witted in deceit, But only as the Crane here from the Crab!” When the Teacher had finished this discourse, showing that “Not now only, O mendicants, has this man been outwitted by the country robe-maker, long ago he was outwitted in the same way,” he established the connexion, and summed up the Jātaka, by saying, “At that time he was the Jetavana robe-maker, the crab was the country robe-maker, but the Genius of the Tree was I myself.” END OF THE STORY OF THE CRUEL CRANE OUTWITTED.336 No. 39. NANDA JĀTAKA. Nanda on the Buried Gold. “The golden heap, methinks.”—This the Master told while at Jetavana, about a monk living under Sāriputta. He, they say, was meek, and mild of speech, and served the Elder with great devotion. Now on one occasion the Elder had taken leave of the Master, started on a tour, and gone to the mountain country in the south of Magadha. When they had arrived there, the monk
- 76. became proud, followedno longer the word of the Elder; and when he was asked to do a thing, would even become angry with the Elder. The Elder could not understand what it all meant. When his tour was over, he returned again to Jetavana; and from the moment he arrived at the monastery, the monk became as before. This the Elder told the Master, saying— “Lord! there is a mendicant in my division of the Order, who in one place is like a slave bought for a hundred, and in another becomes proud, and refuses with anger to do what he is asked.” Then the Teacher said, “Not only now, Sāriputta, has the monk behaved like that; in a former birth also, when in one place he was like a slave bought for a hundred, and in another was angrily independent.” And at the Elder’s request he told the story. Long ago, when Brahma-datta was reigning in Benāres, the Bodisat came to life again as a landowner. He had a friend, also a landowner, who was old himself, but whose wife was young. She had a son by him; and he said to himself— “As this woman is young, she will, after my death, be taking some husband to herself, and squandering the money I have saved. What, now, if I were to make away with the money under the earth?” And he took a slave in the house named Nanda, went into the forest, buried the treasure in a certain spot of which he informed the slave, and instructed him, saying, “My good Nanda! when I am gone, do you let my son know where the treasure is; and be careful the wood is not sold!”
- 77. Very soon afterhe died; and in due course his son became of age. And his mother said to him “My dear! your father took Nanda the slave with him, and buried his money. You should have it brought back, and put the family estates into order.” And one day he accordingly said to Nanda, “Uncle! is there any money which my father buried?” “Yes, Sir!” said he. “Where is it buried?” “In the forest, Sir.” “Then come along there.” And taking a spade and a bag, he went to the place whereabouts the treasure was, and said, “Now, uncle, where is the money?” But when Nanda had got up on to the spot above the treasure, he became so proud of it, that he abused his young master roundly, saying, “You servant! You son of a slave-girl! Where, then, did you get treasure from here?” The young master made as though he had not heard the abuse; and simply saying, “Come along, then,” took him back again. But two or three days after he went to the spot again; when Nanda, however, abused him as before. The young man gave him no harsh word in reply, but turned back, saying to himself,— “This slave goes to the place fully intending to point out the treasure; but as soon as he gets there, he begins to be insolent. I don’t understand the reason of this. But there’s that squire, my father’s friend. I’ll ask him about it, and find out what it is.” So he went to the Bodisat, told him the whole matter, and asked him the reason of it.
- 78. Then said theBodisat, “On the very spot, my young friend, where Nanda stands when he is insolent, there must your father’s treasure be. So as soon as Nanda begins to abuse you, you should answer, ‘Come now, slave, who is it you’re talking to?’ drag him down, take the spade, dig into that spot, take out the treasure, and then make the slave lift it up and carry it home!” And so saying he uttered this verse— “The golden heap, methinks, the jewelled gold, Is just where Nanda, the base-born, the slave, Thunders out swelling words of vanity!” Then the young squire took leave of the Bodisat, went home, took Nanda with him to the place where the treasure was, acted exactly as he had been told, brought back the treasure, put the family estates into order; and following the exhortations of the Bodisat, gave gifts, and did other good works, and at the end of his life passed away according to his deeds. When the Teacher had finished this discourse, showing how formerly also he had behaved the same, he established the connexion, and summed up the Jātaka, “At that time Nanda was the monk under Sāriputta, but the wise squire was I myself.” END OF THE STORY OF NANDA ON THE BURIED GOLD.337 No. 40. KHADIRANGĀRA JĀTAKA. The Fiery Furnace.
- 79. “Far rather willI fall into this hell.”—This the Master told while at Jetavana, about Anātha Piṇḍika. For Anātha Piṇḍika having squandered fifty-four thousands of thousands in money on the Buddhist Faith about the Monastery, and holding nothing elsewhere in the light of a treasure, save only the Three Treasures (the Buddha, the Truth, and the Order), used to go day after day to take part in the Three Great Services, once in the morning, once after breakfast, and once in the evening. There are intermediate services too. And he never went empty- handed, lest the lads, and the younger brethren, should look to see what he might have brought. When he went in the morning he would take porridge; after breakfast ghee, butter, honey, molasses, and so on; in the evening perfumes, garlands, and robes. Thus offering day after day, the sum of his gifts was beyond all measure. Traders, too, left writings with him, and took money on loan from him up to eighteen thousands of thousands, and the great merchant asked it not again of them. Other eighteen thousands of thousands, the property of his family, was put away and buried in the river bank; and when the bank was broken in by a storm they were washed away to the sea, and the brazen pots rolled just as they were—closed and sealed—to the bottom of the ocean. In his house again a constant supply of rice was ordered to be kept in readiness for five hundred members of the Order, so that the Merchant’s house was to the Order like a public pool dug where four high roads meet; and he stood to them in the place of father and mother. On that account even the Supreme Buddha himself used to go to his residence; and the Eighty Chief Elders also; and the number of other monks coming and going was beyond measure. Now his mansion was seven stories high, and there were seven great gates to it, with battlemented turrets over them; and in the fourth turret there dwelt a fairy who was a heretic. When the Supreme Buddha entered the house, she was unable to stop up above in the turret, but used to bring her children downstairs and
- 80. stand on theground floor; and so she did when the Eighty Chief Elders, or the other monks were coming in or going out.338 And she thought, “So long as this mendicant Gotama and his disciples come to the house, there is no peace for me. I can’t be eternally going downstairs again and again, to stand on the ground floor; I must manage so that they come no more to the house.” So one day, as soon as the chief business manager had retired to rest, she went to him, and stood before him in visible shape. “Who’s there?” said he. “It’s I; the Fairy who dwells in the turret over the fourth gate.” “What are you come for?” “You are not looking after the Merchant’s affairs. Paying no thought to his last days, he takes out all his money, and makes the mendicant Gotama full of it. He undertakes no business, and sets no work on foot. Do you speak to the Merchant so that he may attend to his business; and make arrangements so that that mendicant Gotama and his disciples shall no longer come to the place.” But the other said to her, “O foolish Fairy! the Merchant in spending his money spends it on the religion of the Buddhas, which leadeth to salvation. Though I should be seized by the hair, and sold for a slave, I will say no such thing. Begone with you!” Another day the Fairy went to the Merchant’s eldest son, and persuaded him in the same manner. But he refused her as before. And to the Merchant himself she did not dare to speak. Now by constantly giving gifts, and doing no business, the Merchant’s income grew less and less, and his wealth went to ruin. And as he sank more and more into poverty, his property, and his dress, and his furniture, and his food were no longer as they had been. He nevertheless still used to give gifts to the Order; but he was no longer able to give of the best.
- 81. One day whenhe had taken his seat, after saluting the Teacher, he said to him, “Well, householder! are gifts still given at your house?” “They are still being given, Lord,” said he, “but only a mere trifle of stale second day’s porridge.” Then said the Master to him, “Don’t let your heart be troubled, householder, that you give only what is unpleasant to the taste. For if the heart be only right, a gift given to Buddhas, or Pacceka Buddhas,339 or their disciples, can never be otherwise than right. And why? Through the greatness of the result. For that he who can cleanse his heart can never give unclean gifts is declared in the passage— If only there be a believing heart, There is no such thing as a trifling gift To the Mortal One, Buddha, or his disciples. There is no such thing as a trifling service To the Buddhas, to the Illustrious Ones; If you only can see the fruit that may follow, E’en a gift of stale gruel, dried up, without salt! And again he said to him, “Householder! although the gift you are giving is but poor, you are giving it to the Eight Noble Beings.340 Now when I was Velāma, and gave away the Seven Treasures, ransacking the whole continent of India to find them, and kept up a great donation, as if I had turned the five great rivers into one great mass of water, yet I attained not even to taking refuge in the Three Gems, or to keeping the Five Precepts, so unfit were they who received the gifts. Let not your heart be troubled, therefore, because your gifts are trifling.” And so saying, he preached to him the Velāmika Sutta. Now the Fairy, who before had not cared to speak to the Merchant, thinking, “Now that this man has come to poverty, he will listen to what I say,” went at midnight to his chamber, and appeared in visible shape before him.
- 82. “Who’s there?” saidthe Merchant on seeing her. “’Tis I, great Merchant; the Fairy who dwells in the turret over the fourth gate.” “What are you come for?” “Because I wish to give you some advice.” “Speak, then.” “O great Merchant! you take no thought of your last days. You regard not your sons and daughters. You have squandered much wealth on the religion of Gotama the mendicant. By spending your money for so long a time, and by undertaking no fresh business, you have become poor for the sake of the mendicant Gotama. Even so you are not rid of the mendicant Gotama. Up to this very day the mendicants swarm into your house. What you have lost you can never restore again; but henceforth neither go yourself to the mendicant Gotama, nor allow his disciples to enter your house. Turn not back even to behold the mendicant Gotama, but attend to your own business, and to your own merchandize, and so reestablish the family estates.” Then said he to her, “Is this the advice you have to offer me?” “Yes; this is it.” “He whose power is Wisdom has made me immovable by a hundred, or thousand, or even a hundred thousand supernatural beings such as you. For my faith is firm and established like the great mountain Sineru. I have spent my wealth on the Treasure of the Religion that leads to Salvation. What you say is wrong; it is a blow that is given to the Religion of the Buddhas by so wicked a hag as you are, devoid of affection. It is impossible for me to live in the same house with you. Depart quickly from my house, and begone elsewhere!” When she heard the words of the converted, saintly disciple, she dared not stay; and going to the place where she dwelt, she took
- 83. her children bythe hand, and went away. But though she went, she determined, if she could get no other place of abode, to obtain the Merchant’s forgiveness, and return and dwell even there. So she went to the guardian god of the city, and saluted him, and stood respectfully before him. “What are you come here for?” said he. “Sir! I have been speaking thoughtlessly to Anātha Piṇḍika; and he, enraged with me, has driven me out from the place where I dwelt. Take me to him, and persuade him to forgive me, and give me back my dwelling-place.” “What is it you said to him?” “’Henceforth give no support to the Buddha, or to the Order of Mendicants, and forbid the mendicant Gotama the entry into your house.’ This, Sir, is what I said.” “You said wrong. It was a blow aimed at religion. I can’t undertake to go with you to the Merchant!” Getting no help from him, she went to the four Archangels, the guardians of the world. And when she was refused by them in the same manner, she went to Sakka, the King of the Gods, and telling him the whole matter, besought him urgently, saying, “O God! deprived of my dwelling-place, I wander about without a shelter, leading my children by the hand. Let me in your graciousness be given some place where I may dwell!” And he, too, said to her, “You have done wrong! You have aimed a blow at the religion of the Conqueror. It is impossible for me to speak on your behalf to the Merchant. But I can tell you one means by which the Merchant may pardon you.” “It is well, O God. Tell me what that may be!” “People have had eighteen thousands of thousands of money from the Merchant on giving him writings. Now take the form of his
- 84. manager, and withouttelling anybody, take those writings, surround yourself with so many young ogres, go to their houses with the writings in one hand, and a receipt in the other, and stand in the centre of the house and frighten them with your demon power, and say, ‘This is the record of your debt. Our Merchant said nothing to you in byegone days; but now he is fallen into poverty. Pay back the moneys which you had from him.’ Thus, by displaying your demon power, recover all those thousands of gold, and pour them into the Merchant’s empty treasury. There was other wealth of his buried in the bank of the river Aciravatī, which, when the river-bank was broken, was washed away to the sea. Bring that back by your power, and pour it into his treasury. In such and such a place, too, there is another treasure of the sum of eighteen thousands of thousands, which has no owner. That too bring, and pour it into his empty treasury. When you have undergone this punishment of refilling his empty treasury with these fifty-four thousands of thousands, you may ask the Merchant to forgive you.” “Very well, my Lord!” said she; and agreed to what he said, and brought back all the money in the way she was told; and at midnight entered the Merchant’s bed-chamber, and stood before him in visible shape. “Who’s there?” said he. “It is I, great Merchant! the blind and foolish Fairy who used to dwell in the turret over your fourth gate. In my great and dense stupidity, and knowing not the merits of the Buddha, I formerly said something to you; and that fault I beg you to pardon. For according to the word of Sakka, the King of the Gods, I have performed the punishment of filling your empty treasury with fifty-four thousands of thousands I have brought—the eighteen thousands of thousands owing to you which I have recovered, the eighteen thousands of thousands lost in the sea, and eighteen thousands of thousands of owner-less money in such and such a place. The money you spent on the monastery at Jetavana is now all restored. I am in misery so
- 85. long as Iam allowed no place to dwell in. Keep not in your mind the thing I did in my ignorance, but pardon me, O great Merchant!” When he heard what she said, Anātha Piṇḍika thought, “She is a goddess, and she says she has undergone her punishment, and she confesses her sin. The Master shall consider this, and make his goodness known. I will take her before the Supreme Buddha.” And he said to her, “Dear Fairy! if you wish to ask me to pardon you, ask it in the presence of the Buddha!” “Very well. I will do so,” said she. “Take me with you to the Master!” To this he agreed. And when the night was just passing away, he took her, very early in the morning, to the presence of the Master; and told him all that she had done. When the Master heard it, he said “You see, O householder, how the sinful man looks upon sin as pleasant, so long as it bears no fruit; but when its fruit ripens, then he looks upon it as sin. And so the good man looks upon his goodness as sin so long as it bears no fruit; but when its fruit ripens, then he sees its goodness.” And so saying, he uttered the two stanzas in the Scripture Verses: The sinner thinks the sin is good, So long as it hath ripened not; But when the sin has ripened, then The sinner sees that it was sin! The good think goodness is but sin, So long as it hath ripened not; But when the good has ripened, then The good man sees that it was good! And at the conclusion of the verses the Fairy was established in the Fruit of Conversion. And she fell at the wheel-marked feet of the Teacher, and said, “My Lord! lustful, and infidel, and blind as I was, I spake wicked words in my ignorance of your character. Grant me thy pardon!”
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