Note: Descriptions are shown in the official language in which they were submitted.
CA 02772838 2012-03-29
METHODS FOR TREATING AIRCRAFT STRUCTURES
FIELD OF THE INVENTION
The present invention relates to the field of aircraft manufacture and repair;
and also to
preventative maintenance of aircraft. In particular, the invention relates to
the prevention
and repair of structural weaknesses and environmental degradation in aircraft
fuselages and
other structures.
BACKGROUND TO THE INVENTION
The fuselages of many aircraft consist of circumferential frame members,
longitudinal
stringers, and a thin skin, all made from lightweight aluminium. This
construction allows for
a balance of flight properties and weight.
The sheets of aluminium that make up the skin are connected together as lap
joints by
generally two to three rows of rivets. The outer skin later is countersunk at
each rivet
location so the rivet head is flush with the skin, resulting in optimal
aerodynamic properties.
When the skin is subjected to the stresses of normal operation, particularly
in pressurized
commercial aircraft, fatigue damage can occur in the metal sheets and
especially in high
stress locations around fasteners. The problem is exacerbated by the ingress
of
environmental elements and leads to the joint cracking. Crack growth, if left
undetected,
can lead to catastrophic failure, as in the case of Aloha Airlines Flight 243
in 1988. As the
aircraft reached its normal flight altitude of 24,000 feet (7,300 m), a small
section on the left
side of the roof ruptured. The resulting explosive decompression tore off a
large section of
the roof, consisting of the entire top half of the aircraft skin extending
from just behind the
cockpit to the fore-wing area. It was subsequently discovered that the
incident was caused
by the presence of multiple small cracks which arose as a result of
environmental
degradation of the joint located aft of the front port side passenger door.
This phenomena
has subsequently been termed "multi-site damage" (msd)
Since the Aloha incident, aircraft operators have been directed to regularly
check for the
presence of cracks and msd in the fuselage skin. In order to identify the
presence of cracks
before they reach critical lengths, various inspection techniques are
utilized.
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CA 02772838 2012-03-29
While visual inspection is an important part of the detection process, many
naturally
occurring cracks in their initiation are simply too small to see or otherwise
detect. To assist
with the detection of these small and hidden cracks, non-destructive
inspection (NDI)
methods are used. NDI methods can also be used to detect cracks that exist
under paint and
detect areas of corrosion between the layers of skin. Some of the more common
NDI
methods used in aircraft fuselage crack detection are ultrasound and eddy
current methods.
These methods are not capable of detecting all cracks and are particularly
poor in detecting
small naturally occurring defects.
After the Aloha Airlines Flight 243 accident, all 737's with over 50,000
cycles we required to
have their lap joints reinforced with external sheet metal patches. This
modification is
costly, and takes about 15,000 man hours.
Furthermore in April 2011 a fuselage lap joint in a Southwest Airlines Boeing
737-300
aircraft tore an 18 inch hole in the roof, and led to the grounding of 79 of
its older Boeing
737 aircraft for inspections [38, 39]. This resulted in the cancelation of
almost 700 flights
[38, 39]. These inspections, which found cracks in a total of four Southwest
aircraft, [38] led
to the US FAA mandating the inspection of 175 737 aircraft that had
experienced more than
35,000 cycles. There are more than 931 similar aircraft worldwide. The problem
is not
confined to 737 and 727 aircraft. On 26th October 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. This aircraft had experienced less
than 23,000
cycles. This led to the discovery of cracking in other 757 aircraft and a
January 2011 FAA
Airworthiness directive [40] mandating the inspection of all 757-200 and 757-
300 aircraft.
Environmental degradation and subsequent crack initiation and progression is
not just of
importance to commercial airlines. Military aircraft, particularly those with
advanced age,
can also develop environmental degradation and cracking at fastener holes. As
the military
attempts to keep its fleet flight-worthy for longer periods of time,
additional prevention,
inspection, and mitigation methods are being developed to prevent both
environmental
degradation and catastrophic failure.
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T i
When cracks are discovered, they are typically repaired by the application of
external sheet
metal patches. Again, this is a costly and time consuming process. A further
problem is that
the application of patches may actually initiate a weakness in the underlying
structure. Such
undetectected and undetectable cracks can compromise the safety of the
fuselage/wing
skin. These repairs can also locally overstiffen the structure and result in
catastrophic failure
in the fuselage/wing skin as a result of a crack running from repair to
repair.
Externally bonded composite bonded repairs have been developed to address this
problem.
However, these repairs do not prevent the ingress of moisture and hence do not
alleviate
environmental degradation of the structure. Furthermore, to ensure a durable
bond the
structure needs to be heated to approximately 120 C. Additionally, composite
bonded
repairs cannot be used in regions where there is a tight radius of curvature.
It is an aspect of the present invention to overcome or alleviate a problem of
the prior art by
providing a method for preventing or repairing a structural weakness in an
aircraft
structure. A further aspect of the present invention is to overcome or
alleviate a problem of
the prior art by providing a method for preventing environmental degradation
in an aircraft
structure.
The discussion of documents, acts, materials, devices, articles and the like
is included in this
specification solely for the purpose of providing a context for the present
invention. It is not
suggested or represented that any or all of these matters formed part of the
prior art base
or were common general knowledge in the field relevant to the present
invention as it
existed before the priority date of each provisional claim of this
application.
SUMMARY OF THE INVENTION
In a first aspect the present invention provides a method for (i) repairing a
structural
weakness, and/or (ii) preventing or inhibiting the initiation of a structural
weakness, and/or
(iii) preventing or inhibiting the progression of a structural weakness in an
aircraft structure
and/or (iv) preventing the ingress of an environmental element, the method
comprising the
step of bonding a plurality of particles to the structure, the bonding being
effected under
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Y 1
conditions allowing the plurality of metallic particles to form a
substantially continuous
layer.
In one embodiment the method is for (i) repairing a structural weakness,
and/or (ii)
preventing or inhibiting the initiation of a structural weakness, and/or (iii)
preventing or
inhibiting the progression of a structural weakness in an aircraft structure
and (iv)
preventing the ingress of an environmental element.
In a second aspect the present invention provides an aircraft structure
comprising a
substantially continuous metallic layer, the layer being deposited on a
surface of the
structure, the layer being capable of (i) repairing a structural weakness, or
(ii) preventing or
inhibiting the initiation of a structural weakness, or (iii) preventing or
inhibiting the
progression of a structural weakness in the aircraft structure, wherein the
layer comprises a
plurality of metallic particles.
In one embodiment, the layer is capable of (i) repairing a structural
weakness, and/or (ii)
preventing or inhibiting the initiation of a structural weakness, and/or (iii)
preventing or
inhibiting the progression of a structural weakness in an aircraft structure
and (iv)
preventing the ingress of an environmental element.
In one embodiment of the structure the substantially continuous layer is
deposited on the
surface of the aircraft structure by a method as described herein.
In one embodiment of the method or structure, at least a proportion, or
substantially all, of
the particles are metallic particles.
In one embodiment of the method or structure, the bonding does not involve
melting or
fusing of the particles.
In one embodiment of the method or structure, the bonding is achieved by a
cold spray
process, such as supersonic particle deposition.
In one embodiment of the method or structure, the substantially continuous
layer is at least
about 0.05 mm.
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r I
In one embodiment of the method or structure, the substantially continuous
layer has
substantially even depth across the application surface.
In one embodiment of the method or structure, the aircraft structure is a
fuselage
component, such as a sheet metal, and may be a lap joint. Where the structure
is a lap joint
the substantially continuous layer does not extend to cover a junction between
the free end
of a sheet metal component of the lap joint.
In one embodiment, the structural weakness is a crack.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG 1 is a diagram showing the geometry of an edge notched panel, being a
202413 test
specimen. Material is: Aluminium Alloy 2024T3 AlClad 350 mm x 100 mm x 1.27 mm
(0.050").
FIG 2 is a diagram showing the geometry of the panel of FIG 1 with an SPD
doubler.
FIG 3 is a photograph in plan view of a test panel being 2024T3 with SPD
doubler.
FIG 4 is an IR thermal image showing the stresses is an SPD doubler at 11,100
cycles (units in
MPa).
FIG 5 is an IR thermal image showing the stresses is an SPD doubler at 56,100
cycles (units in
MPa).
FIG 6 is a graph showing crack growth histories in SENT tests.
FIG 7 is a diagram showing the location of an SPD strip repair on a 2024T3
test specimen.
FIG 8 is a photograph showing cross-section of an SPD strip.
FIG 9 is an IR thermal image showing the stress field in the skin and the SPD
strip (units in
MPa).
FIG 10 is an IR thermal image showing stresses in an SPD at 3000 cycles (units
in MPa).
FIG 11 is an IR thermal image showing stresses in an SPD at 33,000 cycles
(units in MPa).
FIG 12 is an IR thermal image showing stresses in an SPD at 35,500 cycles
(units in MPa).
FIG 13 is an IR thermal image showing dissipated energy at 33,500 cycles
(units in MPa).
FIG 14 is a graph showing measured crack length histories with and without an
SPD patch.
FIG 15 is a graph showing measured and predicted crack length histories for
the SENT
specimen with an SPD patch.
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r 1
FIG 16 is a diagram showing repair configuration: (a) plan view, (b) cross-
section along
centre line, i.e. x = 0.
FIG 17 shows two images of (a) Cu (bright) on an Al substrate, from [22], and
(b) AlZn onto
an Al substrate, from [23].
FIG 18 is a schematic diagram demonstrating the stresses used to determine K.
FIG 19 is a diagram of a typical finite element mesh of the cracked structure
and the
associated SPD repair. The crack in the base structure (plate) is shown in a
different colour
to the SPD and the remainder of the cracked plate.
FIG 20 is a photograph showing two SPD strips on either side of a 20 mm long
central crack
in a rib stiffened panel.
FIG 21 is a photograph of a delaminated surface of an SPD strip (A), which was
20 mm wide,
showing locations where the fractal dimensions were measured.
FIG 22 is a photograph of a delaminated surface of an SPD strip (B), which was
20 mm wide,
showing locations where the fractal dimensions were measured.
FIG 23A is a cross-sectional view of a lap joint having an SPD doubler.
FIG 23B is an enlarged cross-sectional view of the lap joint of FIG 24A,
showing a single rivet.
FIG 23C is a plan view of the lap joints shown in FIGS 23A and 23B.
FIG 24 is a schematic diagram of a fuselage lap joint specimen, without an SPD
doubler.
FIG 25 is a photograph of a lap joint detailing the SPD application
specification, showing the
region of application. It will be noted that the SPD doubler is applied only
up to the edge of
the upper fuselage skin, and does not extend onto the lower fuselage skin.
Fig 26 is an IR thermal image showing stresses, in MPa, at the critical rows
of fasteners in a
baseline specimen.
Fig 27 is an IR thermal image showing stresses, in MPa, prior to link up.
FIG 28 is an IR thermal image showing stresses, in MPa, in the joint after
approximately
6,500 cycles (Baseline No SPD).
Fig 29 is an IR thermal image showing stresses, in MPa, at approximately
29,000 cycles
(Baseline No SPD).
FIG 30 is an IR thermal image showing stresses in the SPD over the fasteners
at 92,000 cycles
(Test Panel 1).
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FIG 31 is an IR thermal image showing stresses, in MPa, in the lap joint
specimen at 18,000
cycles (Test Panel 2).
FIG 32 is an IR thermal image showing stresses, in MPa, in the lap joint
specimen at 48,000
cycles (Test Panel 2)
DETAILED DESCRIPTION OF THE INVENTION
Reference throughout this specification to "one embodiment" or "an embodiment"
means
that a particular feature, structure or characteristic described in connection
with the
embodiment is included in at least one embodiment of the present invention.
Thus,
appearances of the phrases "in one embodiment" or "in an embodiment" in
various places
throughout this specification are not necessarily all referring to the same
embodiment, but
may. Furthermore, the particular features, structures or characteristics may
be combined in
any suitable manner, as would be apparent to one of ordinary skill in the art
from this
disclosure, in one or more embodiments.
Similarly it should be appreciated that the description of exemplary
embodiments of the
invention, various features of the invention are sometimes grouped together in
a single
embodiment, figure, or description thereof for the purpose of streamlining the
disclosure
and aiding in the understanding of one or more of the various inventive
aspects. This
method of disclosure, however, is not to be interpreted as reflecting an
intention that the
claimed invention requires more features than are expressly recited in each
claim. Rather,
as the following claims reflect, inventive aspects lie in less than all
features of a single
foregoing disclosed embodiment. Thus, the claims following the Detailed
Description are
hereby expressly incorporated into this Detailed Description, with each claim
standing on its
own as a separate embodiment of this invention.
Furthermore, while some embodiments described herein include some but not
other
features included in other embodiments, combinations of features of different
embodiments are meant to be within the scope of the invention, and from
different
embodiments, as would be understood by those in the art.
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For example, in the following claims, any of the claimed embodiments can be
used in any
combination.
In the description provided herein, numerous specific details are set forth.
However, it is
understood that embodiments of the invention may be practiced without these
specific
details. In other instances, well-known methods, structures and techniques
have not been
shown in detail in order not to obscure an understanding of this description.
In the claims below and the description herein, any one of the terms
"comprising",
"comprised of" or "which comprises" is an open term that means including at
least the
elements/features that follow, but not excluding others. Thus, the term
comprising, when
used in the claims, should not be interpreted as being limitative to the means
or elements
or steps listed thereafter. For example, the scope of the expression a method
comprising
step A and step B should not be limited to methods consisting only of methods
A and B. Any
one of the terms "including" or "which includes" or "that includes" as used
herein is also an
open term that also means including at least the elements/features that follow
the term,
but not excluding others. Thus, "including" is synonymous with and means
"comprising".
In a first aspect the present invention provides a method for (i) repairing a
structural
weakness, and/or (ii) preventing or inhibiting the initiation of a structural
weakness, and/or
(iii) preventing or inhibiting the progression of a structural weakness in an
aircraft structure,
and/or (iv) preventing or inhibiting the ingress of an environmental element,
the method
comprising the step of bonding a plurality of particles to the structure, the
bonding being
effected under conditions allowing the plurality of particles to form a
substantially
continuous layer.
In some embodiments of the structure, the method is for (in addition to any
one or all of the
applications (i), (ii) and (iii) recited supra), (iv) preventing or inhibiting
the ingress of an
environmental element. Embodiments having all of the applications (i) to
(iv) are
particularly advantageous given the significant advantages provided in terms
of extended
life of the aircraft structures. Accordingly, a preferred form of the method
provides that the
8
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I
i
method is for (i) repairing a structural weakness, and/or (ii) preventing or
inhibiting the
initiation of a structural weakness, and/or (iii) preventing or inhibiting the
progression of a
structural weakness in an aircraft structure and (iv) preventing the ingress
of an
environmental element.
Applicant has discovered that the deposition of a substantially continuous
layer of particles
over an area of structural weakness (or potential structural weakness) has a
positive effect
on the present or future structural integrity of the aircraft structure. It
has also been shown
that the layer is capable of sealing a substrate against the ingress of
environmental
elements, which in turn may lead to the development of a structural weakness.
The
deposition of a substantially continuous layer has been shown to seal joints,
including
mechanical repairs against the ingress of environmental elements such as
water, salts, air,
acidified rain and the like. Thus, in some embodiments application of the
substantially
continuous layer has a preventative as well as restorative result.
As used herein, the term "repair" is not intended to be construed narrowly to
mean that the
structure must be returned precisely to its original state. It is contemplated
that in some
embodiments, the structure may be returned to a proportion of its original
structural
strength, or indeed or a multiple of its original strength.
The term "prevent" is not intended to be limited to circumstances where the
initiation of a
structural weakness is completely prevented. The initiation of the weakness
may be
delayed in time, or it may manifest as a less severe weakness at initiation.
The term "inhibit" is not intended to mean that progression of the structural
weakness is
completely inhibited. It may just mean that the progression is delayed, or
that it progresses
to an otherwise less severe weakness.
The term "structural weakness" is intended to mean any weakness in the
structure (or in
any part of the structure where the structure is multi-partite) that alters
the ability of the
structure to remain integral upon the application of a force. In the context
of the present
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CA 02772838 2012-03-29
invention a structural weakness includes a crack, a split, a bend, a
deformation, a tear, or
damage occasioned by exposure to an environmental element on an aircraft
structure in the
course of service. It does not include any weakness deliberately inflicted on
a structure, nor
is the term intended to include any alteration in the geometry of a structure
such as may be
occasioned on a gearbox component, or an engine component, for example.
While the step of bonding may involve the melting and/or fusing of the
particles (such as
that involved in high velocity or low velocity oxygen fuel thermal spray
coatings), preferred
embodiments of the method do not involve melting or fusing of the particles.
In one
embodiment of the method, the step of bonding the plurality of particles is
accomplished by
bonding directly to the substrate (and also to each other) by the release of
kinetic energy
from the particles. The particles may acquire the kinetic energy by any means,
but the
energy is typically acquired by accelerating the particles to a high velocity
toward the
substrate. Upon impact with the substrate, the particles deform (typically
flattening) and
form a consolidated structure up to several centimetres thick.
Impact of the panicles with a solid surface at sufficient velocity causes
plastic deformation
and bonding with the underlying material without the creation of heat affected
zones which
are typical of other deposition processes and which are undesirable in many
structural
applications. Bonding is a result of high strain rate deformation and
adiabatic shear
instabilities and the bond interface. Specific advantages of this technology
include but are
not limited to the following
a. SPD produces an excellent bond with the substrate
b. SPD can be used to create almost any required thickness.
c. SPD produces coatings with very few defects Reusable for reclamation
of eroded surfaces and application of wear resistant coating. (SPD enables
the continuing reuse of the base material).
d. Can be applied to recover damaged geometry without adversely affecting
the
substrate (no distortion, heat affected zones or embrittlement).
CA 02772838 2012-03-29
. 1
e. Initial trials have shown that it can be used to enhance structural
integrity
through localized strengthening which may reduce the initiation of cracks or
propagation of cracks.
f. SPD can be used to seal a joint and a riveted repair against the ingress
of
environmental elements thereby alleviating the environmental degradation
of the structure.
8. Significant reduction in occupational health and safety risks
associated with a
number of current in-service applied coatings (e.g. cadmium and hexavalent-
chromium-containing compounds)
Such embodiments of the method are operative at temperatures below the melting
point of
the particle used in the method. In some embodiments, the method is operative
at a
temperature of less than 90%, 80%, 70%, 60%, 50%, 40%, 30%, 20%, 10%, 9%, 8%,
7%, 6%,
5%, 4%, 3%, 2% or 1% of the melting point of the particle.
One particularly useful method for bonding particles (and particularly
metallic particles) to
the aircraft structure is a cold spray method. Cold spray methods are known in
the art, and
are characterized by the application of particles to a substrate at
temperatures below the
melting point of the particles. As used herein, the term "cold spray" is
intended to include
any coating process utilizing a high-speed gas jet to accelerate a plurality
of particles toward
a substrate whereby the particles consolidate on impact. In many cases, the
process is
conducted at a temperature that is substantially lower than the particle
melting point.
The term 'cold spray' has been used to describe this process due to the
relatively low
temperatures of the expanded gas stream that exits the spray nozzle. After
exiting the
nozzle, the particles are impacted onto a substrate, where the solid particles
deform and
create a bond with the substrate. As the process continues, particles continue
to impact and
form bonds with the previously consolidated material resulting in a uniform
deposit with
very little porosity and high bond strength.
Since bonding of the powder to the substrate, as well as the cohesion of the
deposited
material, is accomplished in the solid state at low temperatures, the
characteristics of the
cold sprayed material is advantageous in the context of the present invention.
Because
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CA 02772838 2012-03-29
particle oxidation as well as deleterious tensile stresses that occur during
thermal
contraction are minimized, the cold spray process has the ability to produce
materials with
comparatively superior bond strength to the substrate and greater cohesive
strength.
The Examples herein demonstrate that fatigue performances of cracked metallic
structures
having a cold spray applied metallic layer under constant amplitude loading
are significantly
improved over untreated structures. The experimental data demonstrates that
the baseline
specimens accrued damage more rapidly and that crack growth was significantly
greater
than the corresponding treated panels. In the majority of tests cases the SPD-
treated panels
showed little evidence of damage/crack growth. A prediction of the fatigue
performance of
a treated single edge notch coupon was made using SIF values calculated via an
approximate analysis and the resultant crack length history is in agreement
with
experimental data. Weight function solutions for repairs to centre cracked
panels have also
been developed and validated via three dimensional finite element analysis.
In some embodiments of the method the substantially continuous layer is formed
by
exposing the structure to a high velocity (typically between 300 and 1200 m/s)
stream of
solid-phase particles, which have been accelerated by a supersonic gas flow,
typically
nitrogen or helium, at a temperature that can range between about 400 and 900
C.
Cold spray processes are known by alternative names including supersonic
particle
Deposition (SPD); dynamic spray high velocity powder deposition, kinetic
spraying, and
Kinetic Energy Metallization.
The present invention is a significant departure from the accepted uses of
cold spray
technology. Previously, the method has been used as a coating, much like
paint, or to
restore geometries in worn parts. The use of the process to result in an
aircraft structures
having improved structural characteristics is an advance in the art, providing
economic and
safety advantages.
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In one embodiment, the cold spray process is a low pressure cold spray process
comprising
injecting the plurality of particles in the diverging section of the spray
nozzle from a low-
pressure gas supply.
In low-pressure cold spray, air or nitrogen at relatively low pressure-80-140
psi¨is also
preheated, up to 550 C, then forced through a DeLaval nozzle. At the diverging
side of the
nozzle, the heated gas is accelerated to about 600 m/s. Powder feedstock is
introduced
downstream in the diverging section and accelerated toward the substrate. As
the
applicability of cold spray technology expands to new and unique areas of
application, there
has been an increasing number of commercially available, ready-to-use cold
spray systems
introduced into the marketplace.
In high-pressure cold spray, helium or nitrogen at high pressure, up to 1,000
psi, is
preheated--up to 1,000 C--and then forced through a converging-diverging
DeLaval nozzle.
At the nozzle, the expansion of the gas produces the conversion of enthalpy
into kinetic
energy, which accelerates the gas flow to supersonic regime-1,000 m/s¨while
reducing its
temperature. The powder feedstock is introduced axially into the gas stream,
prior to the
nozzle throat. The accelerated solid particles impact the substrate with
enough kinetic
energy to induce mechanical and/or metallurgical bonding.
The skilled artisan appreciates that both high pressure and low pressure cold
spray
processes may be operable in the context of the present invention.
Of greater relevance than the pressure per se is the velocity at which the
particles are
propelled toward the substrate. Pressure is one parameter that will influence
velocity,
however other factors such as particle size and particle weight will have an
effect. Typically,
the process is operated such that the particles are propelled at or exceeding
a minimum
velocity that is sufficient to provide adhesion of particles, and/or provide
an acceptable
porosity in the resultant coating, and/or provide an acceptable deposition
efficiency.
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However, the velocity should not be so high as to damage the substrate, or
result in the
deflection of significant amounts of particle off the substrate or the
building particulate
layer.
Preferred velocities for a given application vary according to the powder
type. For some
powders a low pressure unit will generate a sufficient velocity to achieve the
required
adhesion, porosity or deposition efficiency. For aluminium powder (which is a
preferred
species of particle in the present methods) a low pressure system may achieve
the desired
outcome if operated at its upper limits. However, it is more typical for a
high pressure unit
to be used in the present methods.
The skilled artisan is enabled to adapt a cold spray method to be operable in
the context of
the present methods. For example, where a particular hardness is required in
the
consolidated metal layer relatively simple particle impact models as applied
to empirical
models for flow stress and hardness may result in useful predications of the
hardness
resulting from a cold spray metallic powder deposition. In particular,
reference is made to
the modelling of Champagne et al (Modelling Simul. Mater. Sc!. Eng. 18 (2010)
065011
(8pp)). These authors show that strain hardening of cold sprayed deposits is a
result of the
flattening of the particles as they impact and bond with the surface. Particle
impact velocity
is the principle controlled parameter of the cold spray process, where
particle velocity and
material properties determine particle flattening.
A constitutive model often used for high strain rate deformation is that of
Johnson and Cook
[5]. This model includes strain hardening, strain rate hardening and thermal
softening
effects during deformation.
A number of variables may be routinely manipulated to achieve a desired
outcome for a
particular application. While not all variables must necessarily be considered
to achieve a
desired outcome, some of which may be considered include the following non-
limiting
parameters:
a. Substrate material
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i. Type
Condition
Surface Finish
b. Powder
i. Material Type
Material Condition
Size
iv. Shape
c. Application Nozzle
Material Type
Nozzle Shape
d. Carrier Gas
i. Type
e. Deposition Parameters
I. Gas input pressure
Gas expansion temperature
iii. Deposition flow rates
iv. Deposition transition rates
The particles may compromise a single species of particle, or multiple
species. The plurality
of particles may be metallic particles, polymer particles or composite
particles. For aircraft-
related application the particles are typically metallic particles fabricated
from any
elementary or alloyed metal, including (but not limited to) aluminium, zinc,
tin, copper,
nickel, titanium, tantalum, cobalt, iron, niobium, molybdenum, and tungsten.
Preferably
the metallic particles are aluminium particles. The particles are typically
utilized in the form
of a commercially available powder, generally ranging in size from about 5 to
about 100 m.
The choice of particle is within the ability of the skilled artisan.
Generally, the particle is
composed of the same material as the substrate.
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CA 02772838 2012-03-29
. .
In one embodiment of the method, the substantially continuous layer is
substantially dense
and/or substantially impervious to a liquid, including water, any polar
solvent or any
nonpolar solvent. An advantage of such layers is that weather is excluded from
any
underlying surface that would normally be prone to corrosion, this enhancing
the operation
life of the aircraft.
The method may be conducted such that a substantially continuous layer of
particles of any
depth is created. The skilled artisan will be capable of assessing a minimum
required depth
for any given structural result required. For example, where the structural
weakness is
minor or the aircraft structure is not a critical component, a lesser depth
may be
implemented. Conversely, a greater depth may be indicated where the structure
has
significant damage, or where the structure has minor damage but is expected to
be exposed
to high levels of stress during operation.
In some embodiments, the method is adapted to deposit a substantially
continuous layer
having a depth of at least about 0.1 mm, 0.2 mm, 0.3 mm, 0.4 mm, 0.5 mm, 0.6
mm, 0.7
mm, 0.8 mm, 0.9 mm, 1.0 mm, 1.1 mm, 1.2 mm, 1.3 mm, 1.4 mm, 1.5 mm, 1.6 mm,
1.7 mm,
1.8 mm, 1.9 mm. 2.0 mm, 2.1 mm, 2.2 mm, 2.3 mm, 2.4 mm, 2.5 mm, 2.6 mm, 2.7
mm, 2.8
mm, 2.9 mm, 3.0 mm, 3.1 mm, 3.2 mm, 3.3 mm, 3.4 mm, 3.5 mm, 3.6 mm, 3.7 mm,
3.8 mm,
3.9 mm, 4.0 mm, 4.1 mm, 4.2 mm, 4.3 mm, 4.4 mm, 4.5 mm, 4.6 mm, 4.7 mm, 4.8
mm, 4.9
mm, 5.0 mm, 5.5 mm, 6.0 mm, 6.5 mm, 7.0 mm, 7.5 mm, 8.0 mm, 8.5 mm, 9.0 mm,
9.5 mm,
10 mm, 11 mm, 12 mm, 13 mm, 14 mm, 15 mm, 16 mm, 17 mm, 18 mm, 19 mm, 20 mm,
21
mm, 22 mm, 23 mm, 24 mm, 25 mm, 26 mm, 27 mm, 28 mm, 29 mm, 30 mm. Preferably,
the substantially continuous layer has a depth of at least about 0.05 mm. At
this depth, a
layer deposited by SPD may be capable of achieving one of (i) repairing a
structural
weakness, (ii) preventing or inhibiting the initiation of a structural
weakness, (iii) preventing
or inhibiting the progression of a structural weakness in an aircraft
structure and (iv)
preventing the ingress of an environmental element the aims. For typical
applications in
aircraft, the substantially continuous layer has a depth of from about 0.2 mm
to about 4
mm. Depths of at least about 0.2 mm have greater utility in structural aspects
of the
invention.
16
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CA 02772838 2012-03-29
. ,
In some embodiments, the method is adapted to deposit a substantially
continuous layer
having a depth of at most about 0.1 mm, 0.2 mm, 0.3 mm, 0.4 mm, 0.5 mm, 0.6
mm, 0.7
mm, 0.8 mm, 0.9 mm, 1.0 mm, 1.1 mm, 1.2 mm, 1.3 mm, 1.4 mm, 1.5 mm, 1.6 mm,
1.7 mm,
1.8 mm, 1.9 mm. 2.0 mm, 2.1 mm, 2.2 mm, 2.3 mm, 2.4 mm, 2.5 mm, 2.6 mm, 2.7
mm, 2.8
mm, 2.9 mm, 3.0 mm, 3.1 mm, 3.2 mm, 3.3 mm, 3.4 mm, 3.5 mm, 3.6 mm, 3.7 mm,
3.8 mm,
3.9 mm, 4.0 mm, 4.1 mm, 4.2 mm, 4.3 mm, 4.4 mm, 4.5 mm, 4.6 mm, 4.7 mm, 4.8
mm, 4.9
mm, 5.0 mm, 5.5 mm, 6.0 mm, 6.5 mm, 7.0 mm, 7.5 mm, 8.0 mm, 8.5 mm, 9.0 mm,
9.5 mm,
mm, 11 mm, 12 mm, 13 mm, 14 mm, 15 mm, 16 mm, 17 mm, 18 mm, 19 mm, 20 mm, 21
10 mm, 22 mm, 23 mm, 24 mm, 25 mm, 26 mm, 27 mm, 28 mm, 29 mm, or 30 mm.
It will be understood that depth of the layer may be adjusted by building up
layers of
particles in the course of the method. In some embodiments, the substantially
continuous
layer has substantially even depth across the application surface.
In one embodiment of the method, the aircraft structure comprises a single
part. The
component may be any part of an aircraft for which structural integrity is
important, such as
a fuselage component. In a preferred embodiment of the method the component is
a sheet
metal forming the "skin" of the fuselage; an engine cowling, or a flight
control such as a
wing (including a flap, aileron, spoiler or winglet), tail (including a
rudder, elevator or
stabilizer), Sheet metal components are thin, relatively flexible and are
particularly prone to
fatigue and cracking. The cold spray application of aluminium particles to the
sheet metal
can increase the life of the part, and also repair any structural weakness
before it leads to a
catastrophic structural failure.
In one embodiment, the aircraft structure has two or more components. The
present
methods have been found to be particularly advantageous for the treatment of
lap joints,
and particularly riveted lap joints. Typically, such joints comprise an upper
skin which
overlaps an inner skin, with the overlapping area being secured by rows of
fasteners, such as
rivets. Often, three rows of rivets are provided across the overlapping area.
17
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CA 02772838 2012-03-29
In one embodiment the aircraft structure is a joint. The joint may comprises
one or more
fastener holes, such as those found in a lap joint. In that case, the method
comprises the
step of bonding a plurality of particles on, over or about the one or more
fastener holes, the
bonding being effected under conditions allowing the plurality of particles to
form a
substantially continuous layer
In some embodiments, the fastener holes include a fastener. In that case the
method
comprises the step of bonding a plurality of particles on, over or about the
one or more
fasteners, the bonding being effected under conditions allowing the plurality
of particles to
form a substantially continuous layer.
Applicant proposes that for structural joints (such as lap joints) and also
riveted repairs,
environmental elements are capable of entering the skin splice or joint during
service. This
leads to degradative effects between the two mating surfaces and also around
fasteners. In
the case of fuselage lap joints the load in the upper skin is transmitted to
the lower skin
through the rivets. These rivets have countersunk heads resulting in what is
termed "knife
edges" in the upper skin. The aggressive environment coupled with the high
stresses at
these knife edges results in crack initiation. It is this crack initiation
(which is exacerbated by
environmental degradation) that results in crack growth at the fuselage joint.
Applicant further proposes that by both sealing the fasteners and reducing
stresses in the
upper skin, the problem of corrosion initiated cracking at fuselage joints is
ameliorated or
overcome. In addition the life of the joints is also increased. This may be
achieved by the
supersonic particle deposition (SPD) of a substantially continuous layer over
the riveted
region lap joint, riveted repairs, and other structural joints in thin skinned
aircraft structure.
Turning now to Fig 23A there is shown diagrammatically and in cross-section a
lap joint
comprised of an upper sheet metal component 1, a lower sheet metal component
2, a rivet
hole 4 passing through components 1 and 2, with a rivet 3. An SPD layer 5 has
been applied
such that the metal particles are bonded to the sheet metal component 1, the
rivet 3, and
an upper region of the rivet hole 4. It will be noted that edges of the layer
5 are bevelled,
18
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CA 02772838 2012-03-29
. .
and that the layer 5 does not extend to cover the junction between the free
end of sheet
metal component 1 and the underlying second sheet metal component 2.
Fig 23B shows an enlarged view of the lap joint of Fig 23, better showing the
countersinking
of the rivet 3 within the joint.
A plan view of the lap joint is shown in Fig 23C showing the three rows of
rivets. The
overlying SPD layer covers the area defined by 5. The edge of the application
area 5
coincides with the edge of the upper sheet metal component 1 such that the SPD
layer does
not extend onto underlying sheet metal component 2, nor does the deposited
material
enter the interface between sheet metal components 1 and 2.
In one embodiment of the invention the structural weakness is a crack.
Preferably the crack
is initiated at the periphery of an aperture in the aircraft structure.
Preferably the aperture
is a fastener hole. Preferably the fastener hold is a faster hole adapted to
receive a faster
having a countersunk head. In one embodiment the crack is one formed by a
force exerted
by the fastener against an edge of a fastener hole. The initiation of the
crack may be due to
normal fatigue, and may aggravated by the presence of damage as a result of
environmental
ingress to the location
In another aspect the present invention provides an aircraft structure
comprising a
substantially continuous layer, the layer being deposited on a surface of the
structure, the
layer being capable of (i) repairing a structural weakness, and/or (ii)
preventing or inhibiting
the initiation of a structural weakness, and/or (iii) preventing or inhibiting
the progression of
a structural weakness in the aircraft structure, and/or (iv) preventing the
ingress of an
environmental element, wherein the layer comprises a plurality of particles.
In some embodiments of the structure, the layer is capable (in addition to any
one or all of
the capabilities (i), (ii) and (iii) recited supra) of (iv) preventing or
inhibiting the ingress of an
environmental element. Embodiments having all of the capabilities (i) to (iv)
are particularly
advantageous with regards to the operable life of an aircraft structure.
Accordingly, a
19
CA 02772838 2012-03-29
preferred form of the method provides that the layer is capable of (i)
repairing a structural
weakness, and/or (ii) preventing or inhibiting the initiation of a structural
weakness, and/or
(iii) preventing or inhibiting the progression of a structural weakness in an
aircraft structure
and (iv) preventing the ingress of an environmental element.
In some embodiments, the substantially continuous layer has a depth of at
least about 0.1
mm, 0.2 mm, 0.3 mm, 0.4 mm, 0.5 mm, 0.6 mm, 0.7 mm, 0.8 mm, 0.9 mm, 1.0 mm,
1.1 mm,
1.2 mm, 1.3 mm, 1.4 mm, 1.5 mm, 1.6 mm, 1.7 mm, 1.8 mm, 1.9 mm. 2.0 mm, 2.1
mm, 2.2
mm, 2.3 mm, 2.4 mm, 2.5 mm, 2.6 mm, 2.7 mm, 2.8 mm, 2.9 mm, 3.0 mm, 3.1 mm,
3.2 mm,
3.3 mm, 3.4 mm, 3.5 mm, 3.6 mm, 3.7 mm, 3.8 mm, 3.9 mm, 4.0 mm, 4.1 mm, 4.2
mm, 4.3
mm, 4.4 mm, 4.5 mm, 4.6 mm, 4.7 mm, 4.8 mm, 4.9 mm, 5.0 mm, 5.5 mm, 6.0 mm,
6.5 mm,
7.0 mm, 7.5 mm, 8.0 mm, 8.5 mm, 9.0 mm, 9.5 mm, 10 mm, 11 mm, 12 mm, 13 mm, 14
mm, 15 mm, 16 mm, 17 mm, 18 mm, 19 mm, 20 mm, 21 mm, 22 mm, 23 mm, 24 mm, 25
mm, 26 mm, 27 mm, 28 mm, 29 mm, or 30 mm.
In some embodiments, the substantially continuous layer has a depth of at most
about 0.1
mm, 0.2 mm, 0.3 mm, 0.4 mm, 0.5 mm, 0.6 mm, 0.7 mm, 0.8 mm, 0.9 mm, 1.0 mm,
1.1 mm,
1.2 mm, 1.3 mm, 1.4 mm, 1.5 mm, 1.6 mm, 1.7 mm, 1.8 mm, 1.9 mm. 2.0 mm, 2.1
mm, 2.2
mm, 2.3 mm, 2.4 mm, 2.5 mm, 2.6 mm, 2.7 mm, 2.8 mm, 2.9 mm, 3.0 mm, 3.1 mm,
3.2 mm,
3.3 mm, 3.4 mm, 3.5 mm, 3.6 mm, 3.7 mm, 3.8 mm, 3.9 mm, 4.0 mm, 4.1 mm, 4.2
mm, 4.3
mm, 4.4 mm, 4.5 mm, 4.6 mm, 4.7 mm, 4.8 mm, 4.9 mm, 5.0 mm, 5.5 mm, 6.0 mm,
6.5 mm,
7.0 mm, 7.5 mm, 8.0 mm, 8.5 mm, 9.0 mm, 9.5 mm, 10 mm, 11 mm, 12 mm, 13 mm, 14
mm, 15 mm, 16 mm, 17 mm, 18 mm, 19 mm, 20 mm, 21 mm, 22 mm, 23 mm, 24 mm, 25
mm, 26 mm, 27 mm, 28 mm, 29 mm, or 30 mm.
In some embodiments, the substantially continuous layer has substantially even
depth
across the application surface.
In one embodiment, the aircraft structure is a single component, such as a
sheet metal. In
One embodiment, the aircraft structure has two or more components, such as a
lap joint.
CA 02772838 2012-03-29
In one embodiment, the substantially continuous layer is formed by a method as
described
herein.
In one embodiment, the aircraft structure is a joint comprising a first
component, and a
second component, the first and second components each having an aperture, the
first and
second components jointed by a fastener extending through the apertures, a
substantially
continuous layer over or about at least one of the apertures.
Preferably, the substantially continuous layer does not extend to cover a
junction between
the free end of a first sheet metal component of the lap joint and the face of
a second sheet
metal component of the joint.. This form of the joint is preferred because the
SPD doubler
does not transfer any load between the upper and lower skin. If the joint were
modified in
this way, a substantial amount of validation work would be necessary to fully
characterise
the structural properties of the modified joint.
The present invention will now be more fully described by reference to the
following non-
limiting examples.
EXAMPLE 1: Effect of SPD on the fatigue performance of cracked metallic
structures.
To study the effect of a supersonic particle deposition (SPD) on the fatigue
performance of
cracked metallic structures initial tests were performed on a 350 mm long and
1.27 mm
thick 2024-13 clad aluminium alloy dogbone specimen which contained a
centrally located 2
mm long edge notch, see Figs. 1-3. These initial tests were performed under
constant
amplitude loading with cr. = 181 MPa and R = omin/an. = 0.1. (This stress
level was chosen
since it represents a realistically upper bound on stresses that can be
expected in a thin
wing skin.) Two specimens were tested, one without a deposited metallic layer
(also known
as a "SPD doubler"), and one with a 1 mm thick full width doubler, that
extended over the
working section of the specimen, deposited on either side of the specimen, see
Figs. 2 and
3. The doublers were deposited using a 7075 Aluminium Alloy powder with a
nominal
particle size of between 30 and 40 m
The following deposition parameters were utilized:
21
CA 02772838 2012-03-29
ELAPSED TIME BETWEEN SURAFCE PREPARATION AND COATING: 15-20mins
MAIN GAS PRESSURE (Bar): 40
P/F VESSEL PRESSURE (Bar): 38
TEMPERATURE ( C): 400
P/F HOPPER HEATER: Active
MAIN GAS FLOW (m3/hr): 92-100
P/F GAS QUANTITY (m3/h): 6.5
POWDER FEED RATE (RPM): 2.7rpm RELEASE TEMP( C): 300
PREHEAT TEMP ( C): 350
INCREMENT (mm): -0.25mm
TRAVERSE RATE (mm/s): 250mm/s
STAND OFF (mm): 40mm (where possible)
NUMBER OF LAYERS: 12 ¨ 16 Passes per Patch
DEPOSITION THICKNESS: 0.005 ¨ 0.012"
POST TRETAMENT: None
For the baseline specimen test the crack length was monitored using digital
cameras.
However, whilst there are numerous non-destructive inspection tools that are
commonly
used to monitor crack growth in aircraft structures, i.e. ultrasonics, eddy
currents,
thermography, etc., the present study used Lock-in infra-red thermography to
simultaneously monitor the evolution of the stress and the damage states in
the 2024-T3
skin and the SPD doublers. (At this point it should be noted that to ensure a
uniform
emissivity the surface being monitored was sprayed matt black and that
thermography was
used as a qualitative rather a quantitative measure of the stresses and the
fatigue damage.
Details on the use of lock-in thermography to measure surface stresses and
energy
dissipation are given in [13-15].) The baseline specimen, i.e. without a
doubler, lasted
approximately 35,000 cycles. In contrast the 7075 SPD patched panel test was
stopped after
approximately 60,000 cycles with little, i.e. no evident, damage in the 7075
SPD or crack
growth in the 2024-T3 skin. Figs. 4 and 5 present infrared pictures of the
stress field at
11,100 and 56,100 cycles respectively. These figures show that the stresses in
the SPD
doubler remained essentially unchanged throughout the test.
22
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CA 02772838 2012-03-29
. .
EXAMPLE 2: Single edge notch tension (SENT) SPD strip tests.
To further study the ability of SPD doublers to reduce crack growth tests were
performed on
a single edge notch dogbone specimen, with a geometry as described above and
an (initial)
1.4 mm long edge notch. In the initial base line test there was no SPD and the
specimen was
tested under constant amplitude loading with a peak stress in the working
section of amax =
93.36 MPa and R min irs / 1 = 0.1. This stress level was chosen to
represent a typical
A - , amax,
fuselage skin stress. Crack growth in the 2024-T3 plate was monitored using
digital cameras
and the resultant crack length versus cycles history is shown in Fig. 6.
In the next test, the specimen was first loaded so as to grow a sharp crack.
This first phase of
the test was stopped at 18,886 cycles when the crack length was approximately
3.2 mm. A
10 mm wide and 1 mm thick SPD strip with a nominally (isosceles) triangular
crossection,
see Figs. 7 and 8, was then installed and the test was continued. The Crack
growth in the
2024-13 plate was again monitored using digital cameras whilst the stress
field in both the
SPD strip and the 2024-T3 skin and the degradation in the SPD strip was
monitored using
Lock-in infrared thermography.
An infrared stress image captured shortly after the restart of the test is
shown in Fig. 9. In
this figure the picture was captured at a cyclic stress amplitude Aci, remote
from the centre
line of the specimen, of approximately 53 MPa. This was done so as to not
overly influence
crack growth in the skin. Here it can be seen how the stress field in the SPD
ahead of the
crack is contiguous with that in the plate, i.e. the SPD is taking load in the
region ahead of
the crack.
Hot spots were also noted in the skin outboard of the ends of the SPD strip
which establish
that the SPD strip was indeed pulling load from the skin. This is essential if
the process is to
enhance the damage tolerance of the skin. The resultant crack growth data is
shown in Fig.
6 where it is seen that the use of a 7075 aluminium alloy SPD strip has
significantly reduced
the crack growth rate.
23
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CA 02772838 2012-03-29
A second test was then performed whereby the SPD strip was applied to a 0.3 mm
long
initial edm (electrical discharge machine) crack (notch) where the crack was
not sharpened
(grown) prior to installation of the SPD strip. In this case the test was
stopped after
approximately 345,000 cycles since there was no apparent crack growth at the
notch (crack)
or damage in the SPD.
EXAMPLE 3: Cracking in 7050-T7451 SENT tests.
It is well known that for combat aircraft most of the fatigue life of the
structure is consumed
in the growth of short cracks [16]. Consequently to evaluate the effect of a
SPD repair on
small flaws in aircraft structural components a 3 mm thick SENT (single edge
notch tension)
dogbone specimen, was tested with a Kt = 1.11, with a thin 0.5 mm thick 7075
aluminium
alloy SPD patch on one side. The 7050-T7451 specimen was 350 mm long, 42 mm
wide and
3 mm thick and had a 0.69 mm radius semi-circular edge notch on one side. The
specimen
was tested at a peak stress, in the working section, of 140.0 MPa with R =
0.1. This
corresponds to a peak (remote) load of 17.64 kN with R = 0.1 and was chosen to
represent
the stress, at limit load, in the wing skin of a typical fighter aircraft.
A thin SPD doubler was used so that, in this test program, the damage induced
could be
evaluated, as the crack opened and closed during fatigue loading, in the
interfacial region
between the SPD and the 7050-T7451. This damage could have been reduced by
increasing
the thickness of the SPD thereby lowering the stresses in the underlying 7050-
17451 and
subsequently reducing the opening of the crack. The ability of the SPD doubler
to pull load
from the underlying 7050-T7451 structure is clear from the E-Mode (stress)
Lockin
thermography picture of the stresses on the specimen side with the SPD patch
at 3000
cycles, see Fig. 10. Although the crack in the 7050-T7451 specimen was not
immediately
evident an analysis of the infra-red data associated with the left hand side
of the specimen
shown in this picture, i.e. in the SPD directly over the crack, revealed an
indication of the
crack under the patch. After 33,000 cycles the crack in the 7050-17451 had
grown to a
length of approximately 4.2 mm and the resultant stress picture is shown in
Fig. 11. At this
point there is evidence of delamination damage (disbonding) on the LHS of the
SPD in the
region that lay over the crack.
24
1
CA 02772838 2012-03-29
. =
The extent of the damage is illustrated in Fig. 12 which presents a picture of
the dissipated
energy at 33,500 cycles. (Note that the dissipated energy associated with the
crack tip is
clearly evident in this figure. This is important because it raises the
potential of non-
destructive inspection of the specimen through the SPD doubler.) At 35,500
cycles the crack
had grown to approximately 4.92 mm and the associated stress picture is shown
in Fig. 13
where it can be seen that the delamination in the SPD has grown slightly.
It would thus appear that whilst damage to the SPD interface can result due to
crack
opening the onset of damage does not appear to lead to immediate
(catastrophic) failure in
the SPD. As such damage growth in the SPD can be included in the damage
tolerance
assessment of the SPD repair process. Furthermore, given that there was no
apparent
damage at the ends of the SPD the damage in the central region over the crack
can be
controlled by increasing the thickness of the SPD in this region thereby
reducing the stress in
the 7050-T7541 together with the associated crack opening displacement.
The test was stopped at 37,000 cycles at which stage the crack was
approximately 5.3 mm
long. A plot of the measured crack length versus cycles history is presented
in Fig. 14
together with test data for the case when there was no SPD. Here it will be
noted that the
SPD patch has somewhat reduced the crack growth rate. To further confirm the
ability of
SPD to restore structural integrity and to illustrate the ability to control
the onset of
delamination damage over the crack a test was subsequently performed on a 1 mm
thick
7050-T7451 SENT specimen, with a 0.8 mm long initial edge crack and an in-
plane geometry
as per the previous test. This specimen had two 0.5 mm thick SPD doublers on
either side of
the specimen. The specimen was subjected to a peak (remote) load of 5.88 kN
with R = 0.1
which equates to the same remote stress as in the previous test. In this case
the test was
stopped after 117,000 cycles as there was no apparent crack growth and no
apparent
degradation in the SPD.
EXAMPLE 4: Predicting crack growth in the 7050-t7451 sent test.
This example is directed to prediction of the crack length history seen in the
7050-T7451
SENT test outlined in Section 2.2. Here a 3 mm thick SENT (single edge notch
tension)
dogbone specimen with a thin 0.5 mm thick 7075 aluminium alloy SPD patch on
one side
CA 02772838 2012-03-29
was tested. The specimen was 350 mm long, 42 mm wide and 3 mm thick and had a
0.69
mm radius semi-circular edge notch on one side. The specimen was subjected to
a peak
(remote) load of 17.64 kN with R = 0.1.
The stress intensity factor for a through-the-thickness crack of length c
emanating from the
centre of the notch of radius r is given in [17] as:
K ¨figifa\/(gc) (1)
Where c is the length of the crack emanating from the notch and a is the
stress in the 7050-
T7451 underneath the SPD. The values offi, g4 and fw taken from [17] are:
¨ 1 0.358(p+ 1.425(p2 1.578(p3 +2.156p4 (2)
(I) =1/(1 clr) (3)
g4 K(036 ¨ 0.32/.1(1 c/r) (4)
f,= 1 i 2.7(p2 3.5(p4 + 3.8(p6 (5)
Kt 3.17 (6)
Since the specimen was tested using hydraulic grips the formulae used for fw
was the fixed
displacement expression given in [17]. Let us now attempt to use this solution
to predict
crack growth. Fig. 11-13 revealed that there was (delamination) damage growth
in the SPD
over the crack. Thus as recommended in [18] for composite repairs to cracked
metal skins
the problem was analysed by assuming that the resultant stress intensity
factor was equal
to the solution to the SENT specimen subjected to a stress field ao which
corresponds to the
stress in the (base) specimen under the SPD in the absence of a crack. The
DSTO Combat
and Trainer Aircraft Group [19-21] have shown that the growth of small flaws
in 7050-
17451 conforms to the Generalised Frost-Dugdale crack growth law, viz:
da (IN ¨ C (AK K mo,,P/ 6-07 (1 Kmax 11<c (7)
26
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CA 02772838 2012-03-29
where C., y and lc are material constants and a, K. and AK are crack lengths,
the maximum
value of the stress intensity factor at cycle N and the range of the stress
intensity factor at
cycle N respectively. The crack length history was predicted by integrating
Eq. (7) using Eqs.
(1)¨(8) with Pmax = 17.4 kN, R = 0.1 and r = 0.69 mm. In this calculation the
values of y, p. C.,
,to be as given in [19,21], viz: y = 3, p = 0.2, C.= 0.50, ay = 460 MPa and,
for this thickness, lc
¨50-65 MPa Vm were taken. (In this analysis a value of Kc = 60 MPa Vm was
used. However,
for this range of loads and crack lengths the value has a small effect on the
crack length
predictions.) The resultant predicted crack length history is shown in Fig. 15
where a good
agreement is seen between the measured and predicted crack length histories.
In this case, as for cracks growing under composite repairs [18] the stress
intensity factor
can be approximated as the solution to the SENT specimen subjected to the
stress field ao
which corresponds to the stress in the (base) specimen under the SPD in the
absence of a
crack. One advantage of this approach is that the computed crack length
history should be
conservative.
EXAMPLE 5: Approximate solutions for centre cracked panels repaired using SPD.
In the previous example the case of a thin (0.5 mm) SPD repair to a small flaw
in a relatively
thick (3 mm) section was considered, and SPD delaminated on either side of the
crack was
noted. In such cases it was reasonable to assume that the dominant effect of
the SPD was to
merely reduce the net section stress [18]. However, for certification purposes
the solution
needed for the stress intensity factor associated with an arbitrary length
crack where SPD
patch is not thin. It is also necessary to establish if, for a given crack
length, the stress
intensity factor range AK is beneath the threshold value A Kth as this will
significantly simplify
the certification process. To this end this example will consider an SPD
repair of thickness tr
to a centre cracked panel, thickness tp, with an interfacial region, thickness
ti, that has been
(potentially) affected by the SPD process subjected to a remote stress a as
shown in Fig. 16.
The SPD process can result in an interfacial region that has been affected by
the SPD process
[22-24]. For the aluminium alloy powders used in SPD repairs the maximum
particle size is
approximately 40 pm. Consequently the thickness (ti) of this region is
generally very small
[22-24] in comparison to the thickness of the underlying plate, i.e. typically
less than 0.1-
0.15 mm, see Fig. 17. As a result this problem is analogous to that of a
bonded repair where
27
CA 02772838 2012-03-29
the interfacial region mimics the adhesive that joins the repair to the plate.
It is known that
for small cracks in metal skins repaired using a composite patch the 2D
solution for the
stress intensity factor is essentially due to the reduction in the stress
field under the repair
whilst for long cracks the stress intensity factor asymptotes to a limiting
stress intensity
factor K¨ as the crack length increases, see [18,25-281 As such it follows
that the 2D
solution for the stress intensity factor associated with small cracks repaired
using SPD is also
essentially due to the reduction in the stress field under the SPD whilst for
long cracks
repaired using SPD the stress intensity factor should also asymptote to a
limiting stress
intensity factor K¨ as the crack length increases. The approximate formulae
for this
asymptote thus follows from [26], see pp. 216-218, viz:
YOI (70 Ora (8)
where
(70 ¨ aEptp/(Eptp I Ertr) (9)
Y is a geometry factor, = 1 for a large centre cracked panel and DI_ is a load
attraction factor
that accounts for the different stiffness of the repaired region. (OL = 1.)
The term nX is given
by the expression
7r,1¨ viEptp//3(1 tpEp/Ertr) (10)
where
(t,/G, tr/3G, f tp/3Gp )/(t1/G1 1 3tr/8G, 1
3t/8G)2 (11)
Here ti, tp and tr are the thickness' of the interface region where, the SPD
has modified the
properties of the plate, of the plate, and of the SPD respectively, G and E
denote the shear
and Young's modulus and the subscripts i, p and r denote their values for the
interfacial
bonding region, the plate, and the SPD repair respectively. (The notation used
in this section
follows that given in [26] pp. 217-218.) This expression, i.e. Eq. (11), is an
extension of the
formulae first developed in [25] in that it allows for the interfacial
thickness ti to be
28
CA 02772838 2012-03-29
negligible. This (allowance) is important since for SPD repairs the modulii of
each region will
generally be comparable and the interfacial thickness ti that is affected by
the SPD process is
expected to be very small. As such the terms in Eq. (11) related to the term
ti/G, are small in
comparison with those terms relating to the SPD repair (tr/3Gr) and the plate
(t/3G).
Consequently the expression for 13 can often be approximated as:
fi (td3Gr 1 tp/3Gp)/(3618G, + 3t/8G)2
(12)
It is expected that, in many instances, the SPD powder, used in the repair,
and the plate
material will have essentially the same modulii, i.e. aluminium plates are
expected to be
repaired using aluminium alloy powders and steel components are likely to be
repaired
using steel powders. In such cases Gr can be approximated by Gp so that Eq.
(12) reduces to:
64Gp/27(t, + tp) (13)
Having established the asymptotic limit it follows from [28-30] that the
functional form of K
as a function of the crack length (a) can be approximated as:
K W(ahr))cro\j(na) (14)
where the function W, viz:
W(a/(n),)). N/Rl +2.23a/(70.))/(1 +301/(7r;.)+ 7(aMEA))2)]
(15)
describes the transition from the small crack solution a40 to the long crack
solution a-> 01,
see [28,30]. Eqs. (14) and (15) reveal that for short cracks the reduction in
the stress
intensity factor is essentially due to the reduction in the stress in the
plate due to the SPD
patch, i.e.
limitKaoµhr- a (16)
For long cracks K tends to its asymptotic limit L. In Eq. (14) the functional
form associated
with [30] has been used rather than that given in [28].
29
CA 02772838 2012-03-29
=
EXAMPLE 6: SPD repairs to cracks in an arbitrary stress field
This example considers the case of an SPD repair to a crack with a total of
length 2a
subjected to an arbitrary stress field. In this instance the solution for the
stress intensity
factor K follows from the above analogy with a composite repair to a crack in
a metal skin
under an arbitrary symmetry stress field [29], viz:
K --W(a/m),)K (17)
where K is the solution to the entre cracked specimen subjected to a stress
ac, which
corresponds to the stress in the (base) specimen under the SPD in the absence
of a crack is
given by
a
K 2 \/(a/m) ay(x)/(a2 x2)1/2dx (18)
0
and cry is the stress in the skin under the SPD in the absence of a crack, see
Fig. 18.
To evaluate the accuracy of this approximation let us consider a 3 mm thick
200 mm 200
mm centre cracked plate repaired using a 3 mm (thick) 200 mm 200 mm SPD patch
subjected to a remote uniform stress (in the skin) of 100 MPa. To this end
three dimensional
finite element models were constructed for: 2, 6, 10, 20, 30, and 40 mm long
cracks. Due to
symmetry considerations only one quarter of the structure needed to be
modelled. In each
case the models had approximately 66,000 three dimensional twenty-one nodded
isoparametric brick elements and approximately 300,000 nodes, see Fig. 19.
There were
eight elements through the thickness of the SPD and eight elements through the
thickness
of the plate. In each case there were ten elements along the crack and the
side length of the
crack tip elements were approximately 1/100th of the length of the crack. The
midside
nodes associated with the near tip elements were moved to the quarter points
so as to
simulate the necessary (1/2 singularity. Bending of the SPD and the plate was
prohibited.
Both the aluminium alloy plate and the SPD were assumed to have a Young's
modulus E =
70,000 MPa and a Poisson's ratio of 0.3. The computed values of the maximum
value of the
CA 02772838 2012-03-29
stress intensity factor Kmax are given in Table 1 below along with the
associated analytical
values, where Eq. (12) was used to compute 13, and the quantity.
Table I
Comparison between predicted and computed stress intensity factors,
a (mm) Finite element Analytical Kmax Upper bound Ku
(MPa Vrn) (MN vm) (MPa vm)
1 4.88 5.03 5.60
3 6.77 6.82 9.71
7.29 7.29 1253
7.56 7.59 17,72
7.60 7.65 21.71
7,60 7.68 25.07
5
Ku or,Orca) (19)
which represents an upper bound on K. Here it can be seen that the stress
intensity factor
associated with SPD repairs does indeed asymptote to a constant value and that
this
10 asymptote is in good agreement with the analytical approximation, i.e.
Eq. (14).
EXAMPLE 7: Quality control assessment tool.
When performing composite repairs to aircraft structural members it is common
practice to
make travelling specimens that are subsequently used to assess the quality of
the repair
15 [11,12]. The challenge is to develop a similar approach for SPD
modifications/repairs to
aircraft structural components. As such this section raises the possibility of
using simple
specimens that are subsequently fatigue tested and the quality of the bond
assessed via the
fractal dimension [31] of the resultant fatigue surfaces. In it this context
it should be noted
that is now known that fracture surfaces can be considered as a fractal set,
see Mandelbrot
20 et al. [31]. In this work Mandelbrot et al. [31] wrote: "When a piece of
metal is fractured
either by tensile or impact loading the facture surface that is formed is
rough and irregular.
Its shape is affected by the metal's microstructure (such as grains,
inclusions, and
precipitates where characteristic length is large relative to the atomic
scale), as well as by
'macrostructural' influences (such as the size, the shape of the specimen, and
the notch
from which the fracture begins). However, repeated observation at various
magnifications
also reveal a variety of additional structures that fall between 'micro' and
'macro' and have
31
i
CA 02772838 2012-03-29
. .
not yet been described satisfactorily in a systematic manner. The experiments
reported
here reveal the existence of broad and clearly distinct zone of intermediate
scales in which
the fracture is modelled very well by a fractal surface." It is also known [32-
34] that, prior to
the onset of rapid fracture, fatigue crack surfaces in metals, that are not
associated with
very small crack lengths, have a fractal box dimension D, as defined in [32],
that lies
between approximately 1.2 and 1. Thus it may be possible to use this
observation to
quantify the quality of the SPD process. To do this travelling specimens would
be fabricated
in parallel with the SPD application. These travelling specimens would
subsequently be
fatigue tested and the associated fractal box dimensions measured. It is
hypothesised that if
D had a value that was near 1.2, or lower, then you would have a process that
produced a
fatigue crack surface that was consistent with that associated with fatigue
crack growth in
the base material, and the process would be acceptable. If it was
significantly greater then it
is hypothesised that the application process may be deficient.
To evaluate this concept the fractal box dimension was measured, the fractal
box
associated with SPD doublers used on a rib stiffened panel deposited using
powders where
there was (subsequently) found to be a quality control issue with the powder,
i.e. it was
found to contain a large proportion of sub 10 micron particles. In this
instance the panels
had two ten mm wide and 200 mm long SPD doublers located on either side of a
centrally
located 20 mm long crack, see Fig. 20.
As a result of the poor quality powder one end of each of the two SPD strips
delaminated
with the locus of the delaminations lying entirely within the SPD, see Figs.
21 and 22. The
fractal box dimensions associated with delamination surfaces on each of the
two SPD strips,
referred to in Figs. 21 and 22 as strips A and B, that delaminated from the
structure were
measured and the resultant values are given in Tables 2 and 3, below.
32
CA 02772838 2012-03-29
Table 2
Fractal box dimension (0) associated with the end of strip A.
Random area 1 within location Random area 2 within location
Lot 1 1.629 1.500
Loc 2 1.409 1.675
Let 3 1.542 1.684
Loc 4 1.416 1.473
Lot-5 1.543 1.530
Lot 6 1.399 1.529
Average 1.49 1.57
Table 3
fractal box dimension (D) associated with the end of strip B.
Random area 1 within location Random area 2 within location
Lot 1 1.673 1.613
Loc 2 1.482 1.521
Loc 3 1.525 1.614
Lot 4 1.551 1.49
Loc 5 1.526 1.516
Lot 6 1.558 1.561
Lac 7 1.578 1.482
Lot 8 1.503 1.593
1.oc 9 1.584 1.563
Average 1.553 1.550
Here it will be noted that in each case the fractal box dimension D was
essentially constant
at each of the locations measured on each of the two delaminated strips.
Furthermore, the
value of the fractal box dimension D was approximately 1.5, see Tables 2 and
3. As such the
fractal box dimension D associated with these two poor quality SPD's differed
significantly
from that associated with macro-scopic fatigue crack growth in metals. Thus
whilst a great
deal more work is needed to validate the hypothesis that D can be used to
quantify the
quality of the SPD it looks to be worthy of further evaluation. It is
interesting to note that
prior to these tests a value of D = 1.5 had only (previously) been found for
very small fatigue
cracks [32,34]. A more detailed discussion of the role of the fractal
dimension D in
describing the nature of the crack tip singularity and in characterising
fatigue crack growth is
given in [20,34-37].
EXAMPLE 8: Application of SPD to an aircraft lap joint.
Specimens have been prepared to evaluate the application of the SPD on a
representative
aircraft lap.
33
CA 02772838 2012-03-29
The specimen geometry was developed as part of the FAA Aging Aircraft Program,
where it
was shown to reproduce the crack length history seen in Boeing 727 and 737
fleet data The
basic specimen used consists of two 2024-T3 clad aluminium alloy sheets 1.016
mm 0.04
inch) thick, fastened with three rows of BACR15CE-5, 1000 shear head counter-
sunk rivets,
3.968 mm (5/32 inch) diameter (Fig 24 ). The width of the specimen was chosen
to coincide
with the typical distance between tear straps of a B-737 aircraft. Since the
amount of out-
of-plane bending in a typical fuselage joint is an important factor in the
fatigue performance
of the joint, the amount of local bending in the specimen was made similar to
that seen in a
typical fuselage joint by testing the specimens bonded back-to-back and
separated by a 25
mm thick honeycomb core. This test configuration was crucial in ensuring that
the
specimens reproduced fleet behaviour, see [40, 41]. As in [40, 41] the upper
row of rivet
holes contained crack initiation sites, induced prior to assembly of the joint
by means of an
electrical spark erosion technique, on either side of the rivet holes. These
initial cracks were
(each) nominally 1.25 mm long. This crack length was chosen so that the
(initial) defect was
obscured by the fastener head and as such was representative of largest
possible
undetectable flaw size.
A 1 mm thick 7075 SPD doubler was deposited over the three rows of fasteners,
(Fig 25 ).
The Powder was deposited utilising the Kinetics 4000 Series CGT equipment with
the Type
33 polycarbonate nozzle. The surface was precleaned utilising 120 Aluminium
Oxide grit at
60 psi. The deposition parameters were as follows:
ELAPSED TIME BETWEEN SURAFCE PREP AND COATING: 15-20mins
MAIN GAS PRESSURE (Bar): 40
P/F VESSEL PRESSURE (Bar): 38
TEMPERATURE ( C): 400
P/F HOPPER HEATER: Active
MAIN GAS FLOW (m3/hr): 92-100 P/F GAS QUANTITY (m3/h): 6.5
POWDER FEED RATE (RPM): 2.7rpm
RELEASE TEMP( C): 300
PREHEAT TEMP ( C): 350
34
CA 02772838 2012-03-29
,
INCREMENT (mm): -0.25mm
TRAVERSE RATE (mm/s): 250mm/s
STAND OFF (mm): 40mm (where possible)
NUMBER OF LAYERS: 12¨ 16 Passes per Patch
DEPOSITION THICKNESS: 0.005 ¨ 0.012"
POST TRETAMENT: None
EXAMPLE 9: Testing of SPD applied to aircraft lap joint
The specimens were tested under constant amplitude loading, with the maximum
and
minimum loads as detailed below.
P Max (kN): 40
PMin (kN) 2
P Mean (kN) 21
Test Frequency (hZ) 5
These loads were determined from operational data obtained for the US DoT MSD
Committee Review Board for the B-737 aircraft, see [40] for more details, and
a stress
picture showing the stresses in the baseline specimens is presented in Figure
26 and a stress
picture just prior to link up of msd is shown in Figure 27. The fatigue
performance of the
baseline (no SPD) specimens is documented in [42].
Here it was found that for specimens without an SPD modification the number of
cycles to
first link up of cracks from adjacent holes occurs at approximately 30,000
cycles. To
illustrate thisand to show the stresses in the baseline joint Figures 28 and
29 present the
stresses in a (baseline) joint at approximately 6,500 and 29,000 cycles
respectively.
For Test Panel 1 the test program revealed that after approximately 110,000
cycles the SPD
doubler was still intact. Furthermore, there was no apparent crack growth at
any of the
fasteners in the lap joint, cracking in the SPD or damage to the bond between
the SPD and
1
CA 02772838 2012-03-29
,
,
the skin/fasteners. This is evident from Figure 30 where we show a close up
view of the
stresses in three rows of rivets at 92,000 cycles.
The test results revealed that the SPD doubler significantly reduces the
stresses in the joint.
This means that the SPD seals the fasteners and continues to do so for more
than 110,000
cycles. This represents a factor more than 3.6 in the Limit of Viability (LOV)
[12] of the joint.
Test Panel 2 also achieved 2 times the LOV (60,000 cycles) even though there
were pre
existing delaminations between the skins (both upper and lower) and the
honeycomb core
(i.e. loss of panel stability) prior to test. Figures 31 and 32 present the
stress distribution in
the SPD Test Panel 2 at 18,000 cycles and 48,000 cycles respectively.
Comparing Figures 26
and 31 a significant reduction in the stresses in the joint is noted. In the
upper section of the
picture 31, an increase in the stresses in the skin at the edge of the SPD
where load is being
attracted up from the skin into the SPD. The stress concentrations in the SPD
over each of
the fasteners are also visible.
From this Example it can be seen that the SPD has remained intact, thereby
ensuring that
the joint is sealed. Although this study has focused on fuselage lap joints
the ability of an
SPD doubler to form a durable bond to both the skin and the fasteners predicts
that this
approach may well be applicable to other problem areas in an aircraft.
36
1
CA 02772838 2012-03-29
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41
