U.S. patent application number 13/922933 was filed with the patent office on 2014-09-04 for air cooled gas turbine components and methods of manufacturing and repairing same.
The applicant listed for this patent is General Electric Company. Invention is credited to Thomas George Holland, David Bruce Patterson, John Howard Starkweather, Thomas John Tomlinson.
Application Number | 20140248425 13/922933 |
Document ID | / |
Family ID | 40084057 |
Filed Date | 2014-09-04 |
United States Patent
Application |
20140248425 |
Kind Code |
A1 |
Patterson; David Bruce ; et
al. |
September 4, 2014 |
AIR COOLED GAS TURBINE COMPONENTS AND METHODS OF MANUFACTURING AND
REPAIRING SAME
Abstract
A method of manufacturing a component suitable for use in a gas
turbine engine, comprising the steps of forming the component from
a substrate having a first surface and a second surface, forming at
least one aperture through the substrate from the first surface to
the second surface having a first open area, applying a first
coating to at least one of the first surface and the second surface
adjacent to the at least one aperture, the aperture remaining at
least partially unobstructed by the first coating, applying a
second coating to the first coating adjacent to the at least one
aperture, the aperture remaining at least partially unobstructed by
the second coating, and removing the second coating from the
aperture, leaving most or all of the first coating to define a
second open area which is smaller than the first open area. In a
further aspect, a method of repairing a component suitable for use
in a gas turbine engine, the method comprising the steps of
removing coatings from the component, repairing any defects in the
substrate of the component, and applying coatings as described
herein.
Inventors: |
Patterson; David Bruce;
(Mason, OH) ; Starkweather; John Howard;
(Cincinnati, OH) ; Holland; Thomas George;
(Dayton, OH) ; Tomlinson; Thomas John; (West
Chester, OH) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Family ID: |
40084057 |
Appl. No.: |
13/922933 |
Filed: |
June 20, 2013 |
Related U.S. Patent Documents
|
|
|
|
|
|
Application
Number |
Filing Date |
Patent Number |
|
|
11926986 |
Oct 29, 2007 |
|
|
|
13922933 |
|
|
|
|
Current U.S.
Class: |
427/142 ;
427/292 |
Current CPC
Class: |
C23C 28/3455 20130101;
C23C 4/02 20130101; Y02T 50/675 20130101; F01D 5/288 20130101; Y10T
428/24322 20150115; Y02T 50/6765 20180501; C23C 28/325 20130101;
F23M 2900/05004 20130101; F01D 9/023 20130101; F23R 3/007 20130101;
C23C 4/01 20160101; F01D 25/12 20130101; Y02T 50/60 20130101; Y02T
50/67 20130101; C23C 28/345 20130101; C23C 28/3215 20130101; F23R
3/06 20130101; F01D 5/005 20130101 |
Class at
Publication: |
427/142 ;
427/292 |
International
Class: |
F01D 25/12 20060101
F01D025/12 |
Claims
1. A method of manufacturing a component suitable for use in a gas
turbine engine, said method comprising the steps of: forming said
component from a substrate having a first surface and a second
surface; forming at least one aperture through said substrate from
said first surface to said second surface, said aperture having a
first open area; applying a first coating to at least one of said
first surface and said second surface adjacent to said at least one
aperture, said aperture remaining at least partially unobstructed
by said first coating; applying a second coating to said first
coating adjacent to said at least one aperture, said aperture
remaining at least partially unobstructed by said second coating;
and removing said second coating from said aperture, leaving most
or all of said first coating to define a second open area which is
smaller than said first open area.
2. The method of claim 1, wherein at least one of said first
coating and said second coating are applied at an angle to said at
least one of said first surface and said second surface.
3. The method of claim 1, wherein at least one of said removal
steps is accomplished by a stream of abrasive media.
4. The method of claim 3, wherein said stream of abrasive media
comprises glass beads suspended in a stream of air.
5. The method of claim 3, wherein said stream of abrasive media is
directed through said at least one aperture from a non-coated side
of said substrate.
6. A method of repairing a component suitable for use in a gas
turbine engine, said component having a substrate with first and
second surfaces and at least one aperture extending through said
substrate from said first surface to said second surface, said
aperture having a first open area, said method comprising the steps
of: removing coatings from said component; repairing any defects in
said substrate of said component; applying a first coating to at
least one of said first surface and said second surface adjacent to
said at least one aperture, said aperture remaining at least
partially unobstructed by said first coating; applying a second
coating to said first coating adjacent to said at least one
aperture, said aperture remaining at least partially unobstructed
by said second coating; and removing said second coating from said
aperture leaving most or all of said first coating to define a
second open area which is smaller than said first open area.
7. The method of claim 6, wherein at least one of said first
coating and said second coating are applied at an angle to said at
least one of said first surface and said second surface.
8. The method of claim 6, wherein at least one of said removal
steps is accomplished by a stream of abrasive media.
9. The method of claim 8, wherein said stream of abrasive media
comprises glass beads suspended in a stream of air.
10. The method of claim 8, wherein said stream of abrasive media is
directed through said at least one aperture from a non-coated side
of said substrate.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application is a divisional of U.S. Ser. No.
11/926,986, filed on Oct. 29, 2007, the entire disclosure of which
is incorporated herein by reference.
BACKGROUND OF THE INVENTION
[0002] The technology described herein relates generally to gas
turbine engines, and more particularly, to air-cooled components
for use in gas turbines and methods of manufacturing and repairing
such components.
[0003] A gas turbine engine includes a compressor for compressing
air which is mixed with a fuel and channeled to a combustor wherein
the mixture is ignited within a combustion chamber for generating
hot combustion gases. At least some known combustors include a dome
assembly, a cowling, and liners to channel the combustion gases to
a turbine, which extracts energy from the combustion gases for
powering the compressor, as well as producing useful work to propel
an aircraft in flight or to power a load, such as an electrical
generator. The liners are coupled to the dome assembly with the
cowling, and extend downstream from the cowling to define the
combustion chamber.
[0004] The operating environment within a gas turbine engine is
both thermally and chemically hostile. Significant advances in high
temperature alloys have been achieved through the formulation of
iron, nickel and cobalt-base superalloys, though components formed
from such alloys often cannot withstand long service exposures if
located in certain sections of a gas turbine engine, such as the
turbine, combustor or augmentor. A common solution is to protect
the surfaces of such components with an environmental coating
system, such as an aluminide coating or a thermal barrier coating
(TBC) system. The latter typically includes an
environmentally-resistant bond coat and a thermal barrier coating
of ceramic deposited on the bond coat. Bond coats are typically
formed from an oxidation-resistant alloy such as MCrAlY where M is
iron, cobalt and/or nickel, or from a diffusion aluminide or
platinum aluminide that forms an oxidation-resistant
intermetallic.
[0005] While thermal barrier coating systems provide significant
thermal protection to the underlying component substrate, internal
cooling of components such as combustor liners is generally
necessary, and may be employed in combination with or in lieu of a
thermal barrier coating. Combustor liners of a gas turbine engine
often require a complex cooling scheme in which cooling air flows
around the combustor and is then discharged into the combustor
through carefully configured cooling holes in the combustor liner.
The performance of a combustor is directly related to the ability
to provide uniform cooling of its surfaces with a limited amount of
cooling air. Consequently, processes by which cooling holes and
their openings are formed and configured are often critical because
the size and shape of each opening determine the amount of air flow
exiting the opening and the distribution of the air flow across the
surface, and affect the overall flow distribution within the
combustor. Other factors, such as local surface temperature of the
liner, are also affected by variations in opening size.
[0006] For combustor liners without a thermal barrier coating,
cooling holes are typically formed by such conventional drilling
techniques as electrical-discharge machining (EDM) and laser
machining. However, EDM cannot be used to form cooling holes in a
combustor liner having a ceramic TBC since the ceramic is
electrically nonconducting, and laser machining is prone to
spalling the brittle ceramic TBC by cracking the interface between
the substrate and the ceramic. Accordingly, cooling holes have been
required to be formed by EDM and/or laser machining prior to
applying the TBC system, limiting the thickness of the TBC which
can be applied or necessitating a final operation to remove ceramic
from the cooling holes in order to reestablish the desired size and
shape of the openings. Conventional processes involve protecting
cooling holes from TBC deposition or complete removal of applied
TBC from the holes to obtain the desired hole geometry. This leaves
the underlying metal surface exposed to hostile environmental
conditions at the hole locations.
[0007] Current repair methods for air-cooled components such as
combustor liners include welding thermal fatigue cracks. The
location of openings in the panels, such as cooling or dilution
holes, and the use of thermal barrier coatings add additional
complexity to the use of welds and patches. In many instances,
protective coatings must be removed from an entire panel and/or an
entire liner to gain access to the underlying metal itself, then
reapplying protective coatings. However, conventional reapplication
processes involve protecting cooling holes from TBC deposition or
complete removal of applied TBC from the holes to obtain the
desired hole geometry. This leaves the underlying metal surface
exposed to hostile environmental conditions at the hole locations.
In some cases, repair of such panels is not a feasible option, and
instead the entire combustor liner is replaced.
[0008] Because conventional designs may rely upon the underlying
metal substrate to define the finished hole geometry in the absence
of a TBC system applied to the hole surfaces, damage to or repair
procedures performed on the holes in the metal substrate may affect
the performance of the repaired part. Accordingly, a method is
desired for manufacturing air-cooled components such as combustor
liners in a manner which is economically and physically feasible,
provides enhanced protection to the substrate in the vicinity of
the cooling holes, and which yields a satisfactory cooling hole
geometry both as-manufactured and as-repaired.
BRIEF SUMMARY OF THE INVENTION
[0009] In one aspect, described herein is a component suitable for
use in a gas turbine engine. The component includes a substrate
defining a surface of the component and has a first surface and a
second surface. At least one aperture extends through the substrate
from the first surface to the second surface, and has a first open
area. The component has a first coating on at least one of the
first surface and the second surface adjacent to the at least one
aperture. The component also has a second coating overlying the
first coating adjacent to the at least one aperture, such that at
least a portion of the first coating is exposed adjacent to the at
least one aperture. The first coating defines a second open area
which is smaller than the first open area.
[0010] In another aspect, described herein is a method of
manufacturing a component suitable for use in a gas turbine engine,
comprising the steps of forming the component from a substrate
having a first surface and a second surface, forming at least one
aperture through the substrate from the first surface to the second
surface having a first open area, applying a first coating to at
least one of the first surface and the second surface adjacent to
the at least one aperture, the aperture remaining at least
partially unobstructed by the first coating, applying a second
coating to the first coating adjacent to the at least one aperture,
the aperture remaining at least partially unobstructed by the
second coating, and removing the second coating from the aperture,
leaving most or all of the first coating to define a second open
area which is smaller than the first open area.
[0011] In a further aspect, described herein is a method of
repairing a component suitable for use in a gas turbine engine, the
component having a substrate with first and second surfaces and at
least one aperture extending through the substrate from the first
surface to the second surface, the aperture having a first open
area, the method comprising the steps of removing coatings from the
component, repairing any defects in the substrate of the component,
applying a first coating to at least one of the first surface and
the second surface adjacent to the at least one aperture, the
aperture remaining at least partially unobstructed by the first
coating, applying a second coating to the first coating adjacent to
the at least one aperture, the aperture remaining at least
partially unobstructed by the second coating, and removing the
second coating from the aperture, leaving most or all of the first
coating to define a second open area which is smaller than the
first open area.
BRIEF DESCRIPTION OF THE DRAWINGS
[0012] The accompanying drawings illustrate several embodiments of
the technology described herein, wherein:
[0013] FIG. 1 is a schematic illustration of an exemplary gas
turbine engine;
[0014] FIG. 2 is a schematic cross-sectional view of an exemplary
combustor assembly that may be used with the gas turbine engine
shown in FIG. 1;
[0015] FIG. 3 is an enlarged perspective view of a portion of an
exemplary combustor liner that may be used with the combustor
assembly shown in FIG. 2;
[0016] FIG. 4 is an enlarged partial cross-sectional view of the
combustor liner shown in FIG. 3 before a coating application;
and
[0017] FIG. 5 is an enlarged partial cross-sectional view of the
combustor liner shown in FIG. 4 after a coating application;
and
[0018] FIG. 6 is an enlarged partial cross-sectional view of the
combustor liner shown in FIG. 5 after removing some coating
material; and
[0019] FIG. 7 is a flowchart illustrating steps associated with an
exemplary manufacturing method; and
[0020] FIG. 8 is a flowchart illustrating steps associated with an
exemplary repair method.
DETAILED DESCRIPTION OF THE INVENTION
[0021] The present invention is generally applicable to air-cooled
components, and particularly those that are protected from a
thermally and chemically hostile environment by a thermal barrier
coating system. Notable examples of such components include the
high and low pressure turbine nozzles and blades, shrouds,
combustor liners and augmentor hardware of gas turbine engines. The
advantages of this invention are particularly applicable to gas
turbine engine components that employ internal cooling and a
thermal barrier coating to maintain the service temperature of the
component at an acceptable level while operating in a thermally
hostile environment.
[0022] FIG. 1 is a schematic illustration of an exemplary gas
turbine engine 10. Engine 10 includes a low pressure compressor 12,
a high pressure compressor 14, and a combustor assembly 16. Engine
10 also includes a high pressure turbine 18, and a low pressure
turbine 20 arranged in a serial, axial flow relationship.
Compressor 12 and turbine 20 are coupled by a first shaft 21, and
compressor 14 and turbine 18 are coupled by a second shaft 22. In
the exemplary embodiment, gas turbine engine 10 is a CFM-56 engine
commercially available from CFM International, Inc., Cincinnati,
Ohio. In another embodiment, gas turbine engine 10 is a CF-34
engine commercially available from GE's Aviation business,
Cincinnati, Ohio.
[0023] In operation, air flows through low pressure compressor 12
and compressed air is supplied from low pressure compressor 12 to
high pressure compressor 14. The highly compressed air is delivered
to combustor 16. Airflow from combustor 16 drives turbines 18 and
20 and exits gas turbine engine 10 through a nozzle (not
numbered).
[0024] FIG. 2 is a schematic cross-sectional view of an exemplary
combustor 16 that may be used with gas turbine engine 10 (shown in
FIG. 1). Combustor 16 includes an outer liner 52 and an inner liner
54 disposed between an outer combustor casing 56 and an inner
combustor casing 58. Outer and inner liners 52 and 54 are spaced
radially from each other such that a combustion chamber 60 is
defined therebetween. Outer liner 52 and outer casing 56 form an
outer passage 62 therebetween, and inner liner 54 and inner casing
58 form an inner passage 64 therebetween. A cowl assembly 66 is
coupled to the upstream ends of outer and inner liners 52 and 54,
respectively. An annular opening 68 formed in cowl assembly 66
enables compressed air entering combustor 16 through a diffuse
opening in a direction generally indicated by arrow A. The
compressed air flows through annular opening 68 to support
combustion and to facilitate cooling liners 52 and 54.
[0025] An annular dome plate 70 extends between, and is coupled to,
outer and inner liners 52 and 54 near their upstream ends. A
plurality of circumferentially spaced swirler assemblies 72 are
coupled to dome plate 70. Each swirler assembly 72 receives
compressed air from opening 68 and fuel from a corresponding fuel
injector 74. Fuel and air are swirled and mixed together by swirler
assemblies 72, and the resulting fuel/air mixture is discharged
into combustion chamber 60. Combustor 16 includes a longitudinal
axis 75 which extends from a forward end 76 to an aft end 78 of
combustor 16. In the exemplary embodiment, combustor 16 is a single
annular combustor. Alternatively, combustor 16 may be any other
combustor, including, but not limited to a double annular
combustor.
[0026] In the exemplary embodiment, outer and inner liners 52 and
54 each include a plurality of overlapped panels 80. More
specifically, in the exemplary embodiment, outer liner 52 includes
five panels 80 and inner liner 54 includes four panels 80. In an
alternative embodiment, both outer and inner liner 52 and 54 may
each include any number of panels 80. Panels 80 define combustion
chamber 60 within combustor 16. Specifically, in the exemplary
embodiment, a pair of first panels 82, positioned upstream, define
a primary combustion zone 84, a pair of second panels 86,
positioned downstream from first panels 82, define an intermediate
combustion zone 88, and a pair of third panels 90, positioned
downstream (direction B in FIG. 3) from second panels 86, and a
pair of fourth panels 92, positioned downstream from third panels
90, define a downstream dilution combustion zone 94.
[0027] Combustor liners may include dilution holes to provide air
into the combustion environment with the combustor, such as to
alter the temperature distribution or combustion characteristics.
Dilution air is introduced primarily into combustor chamber 60
through a plurality of circumferentially spaced dilution holes 96
that extend through either or both of outer and inner liners 52 and
54. In the exemplary embodiment, dilution holes 96 are each
substantially circular. Dilution holes may be adapted (sized,
shaped, and/or arranged) as needed to accomplish the durability and
performance objectives of the particular component and the
particular product application.
[0028] FIG. 3 illustrates an exemplary combustor liner 52 that may
be used with combustor 16. Liner 52 also includes a plurality of
cooling holes 160 formed in the third panel 90 that facilitate
cooling liner 52. Although, only one group of cooling holes 160 is
illustrated in the third panel 90, it should be understood that the
group of cooling holes 160 are spaced circumferentially about the
third panel 90. It should be appreciated that each group of cooling
holes 160 is positioned corresponding hot spots to facilitate
channeling cooling fluid onto the corresponding hot spot. Third
panel 90 includes any number of cooling holes 160 that facilitates
cooling of liner 52.
[0029] During operation of gas turbine engine 10, an inner surface
33 of liner 52 becomes hot and requires cooling. Consequently, in
the exemplary embodiment, cooling features such as cooling holes
160 are positioned in liner 52 to facilitate channeling cooling
fluid onto hot spots of liner 52. More specifically, cooling holes
160 channel cooling fluid from outer passage 62 and/or inner
passage 64 to the combustion chamber 60, thus providing a layer of
cooling fluid to inner surface 33. It should be appreciated that
other embodiments may use any configuration of cooling holes 160
that enables cooling holes 160 to function as described herein.
Similarly, holes 160 could be in liner 54 to cool its outer
surface.
[0030] During operation, as atomized fuel is injecting into
combustion chamber 60 and ignited, heat is generated within
combustion chamber 60. Although air enters combustion chamber 60
through cooling features 160 and forms a thin protective boundary
of air along combustor liner surface 33, a variation in exposure of
combustor liner surfaces to high temperatures may induce thermal
stresses into panels 80. As a result of continued exposure to
thermal stresses, over time, panels 80 may become deteriorated.
[0031] FIG. 4 is an enlarged partial cross-sectional view of a
portion of the combustor liner 52 to illustrate the relationship
between the cooling hole 160 and the liner surface 33, as well as
the axis 220 of the hole 160.
[0032] Referring now to FIGS. 5 and 6, a layer 210 of thermal
barrier material is applied to the combustor liner 52 shown in FIG.
4 on combustor liner surface 33. Thermal barrier material further
insulates combustor liner surface 33 from high temperature
combustion gases. Layer 210 includes an inner layer 212, such as a
bond coat layer, and an outer layer 214, such as a thermal barrier
layer.
[0033] The exemplary methods will be described in terms of an
air-cooled component, such as a combustor liner 52, whose metallic
substrate 33 is protected by a thermal barrier coating system
composed of a bond coat 212 formed on the substrate (inner surface
33), and a ceramic layer 214 adhered to the surface 33 with the
bond coat 212. Bond coat 212 and ceramic layer 214 may each be a
single layer of material, or formed of two or more layers (i.e.,
multi-layer) of appropriate materials. As is the situation with
high temperature components of a gas turbine engine, the surface 33
may be an iron, nickel or cobalt-base superalloy. The bond coat 212
is preferably an oxidation-resistant composition, such as a
diffusion aluminide or MCrAlY, that forms an alumina
(Al.sub.2O.sub.3) layer or scale (not shown) on its surface during
exposure to elevated temperatures. The alumina scale protects the
underlying superalloy surface 33 from oxidation and provides a
surface to which the ceramic layer 214 more tenaciously
adheres.
[0034] The ceramic layer 214 can be deposited by air plasma
spraying (APS), low pressure plasma spraying (LPPS), or physical
vapor deposition (PVD) techniques such as electron beam physical
vapor deposition (EBPVD), the latter of which yields a
strain-tolerant columnar grain structure. An exemplary material for
the ceramic layer 214 is zirconia partially stabilized with yttria
(yttria-stabilized zirconia, or YSZ), though zirconia fully
stabilized with yttria could be used, as well as zirconia
stabilized by other oxides, such as magnesia (MgO), calcia (CaO),
ceria (CeO.sub.2) or scandia (Sc.sub.2O.sub.3).
[0035] The method of this invention entails producing a cooling
hole 160 (shown in FIGS. 4-6) which can project through the ceramic
layer 214, bond coat 212 and surface 33 via an opening 162, to
achieve a configuration for the cooling hole 160 and opening 162
that provides an appropriately metered distribution of cooling air
across the external surface of the component, such as liner 52. As
shown in FIG. 5, the cooling hole opening 162 as initially coated
forms a small opening (having axis 230) aimed at a steep angle
(angle 13) to the surface. As shown in FIG. 6, after removing the
portion of the ceramic layer 214 that is in alignment with the
hole, the opening 162 is at a relatively shallow angle to the
surface 33 such that the cooling air flowing through the opening
162 can be laid down as an effective film over the component
surface during operation.
[0036] FIGS. 7 and 8 illustrate in flow diagram form the exemplary
methods described in greater detail herein. While both methods
share some common steps, the method 200 is particularly suited for
manufacturing of new air-cooled components while the method 300 is
particularly suited for repair and restoration of air-cooled
components during their service life.
[0037] As shown in FIG. 4, a first step of this exemplary method is
to form a hole 160 through the liner 52. A second step is then to
apply as shown in FIG. 5 the bond coat 212 and ceramic layer 214 to
the surface 33. Due to coating buildup at the edges of the hole
160, the resulting hole opening 162 is smaller in cross-sectional
diameter than the cooling hole 160 required for the liner 52, but
is not completely obstructed such that the location of the hole and
at least a portion of its cross-section remain substantially free
of obstruction. For example, for a cooling hole 160 having a
diameter of about 0.035 inch (about 0.9 mm) to about 0.040 inch
(about 1.0 mm), the opening 162 after coating preferably has a
diameter of about 0.020 inch (about 0.5 mm), or roughly half that
intended for the cooling hole 160, such that a "witness hole"
remains visible and accessible through the coatings. Suitable
techniques for forming the hole 160 include EDM, though it is
foreseeable that the hole 160 could be formed by such other methods
as casting, laser, or drilling with an abrasive water jet. As a
result of the drilling operation, the hole 160 has a substantially
uniform circular cross-section, and forms a non-normal angle (angle
a) to the surface 33.
[0038] Once the hole 160 is formed, and the bond coat 212 and
ceramic layer 214 are applied, the component (liner 52) is
processed through a carefully controlled operation that uses a
pressurized fluid stream targeted at the hole 160, such as from the
uncoated side of the liner 52, to produce the cooling hole 160 and
opening 162 shown in FIG. 5. Various fluids could be used, such as
air or water, containing a media such as glass beads or an abrasive
grit to provide an abrasive action on coating materials overlying
the hole 160.
[0039] An operation as described herein has been found to provide
sufficient energy to enlarge the opening 162 to the size desired as
well as the angle desired by removing the ceramic TBC layer but not
the bond coat layer or underlying parent material such as the metal
substrate. Therefore, while the operation removes the ceramic layer
214 most or all of the underlying bond coat 212 remains on the
surface of the opening adjacent to the cooling hole 160, such that
the bond coat layer provides protection for the edges of the liner
in the vicinity of the cooling hole both during manufacture and in
service. Because the operation uses mechanical energy rather than
heat energy, it does not damage or spall the bond coat 212 or
ceramic layer 214 surrounding the hole 160 and forming the edges of
the resulting hole opening 162.
[0040] The method is capable of appropriately sizing and shaping
cooling holes and openings through a ceramic thermal barrier
coating (TBC) and its underlying substrate. The abrasive fluid
stream also serves to finish the hole and its opening, including
the desired size and shape of the hole and opening, without
removing or damaging the ceramic surrounding the cooling hole and
opening.
[0041] If a field returned engine, such as engine 10, indicates
that combustor liner 52 includes at least one deteriorated panel
80, a variety of repair methods may be employed to restore
combustor liner 52 to serviceable condition. These repair methods
may include replacement of the entire liner, a complete panel,
and/or a portion or segment of a liner panel, as well as repair of
cracks such as by welding them closed.
[0042] During a repair operation, all dirt, foreign material, and
coatings are normally removed from a component such as a combustor
liner to permit a detailed inspection of the component. Any defects
in the substrate, such as cracks, are then repaired using suitable
and approved methods such as welding, brazing, or replacement of
discrete sections of the component. Holes such as cooling holes may
be redrilled and/or repaired as needed to restore them to the
appropriate size, shape, and pattern.
[0043] Once the surfaces of the component have been suitably
repaired, protective thermal barrier coatings may be applied to
component surfaces utilizing the exemplary methods described above.
Because the finished opening dimensions are carefully controlled
and are defined by a removable and replaceable coating system as
described herein, it is possible to perform and repeat the repair
process while maintaining finished cooling hole dimensions within
specifications.
[0044] Because components such as deteriorated liners are repaired
using the method described herein, utilizing readily available
coating techniques, combustors may be returned to service using a
repair process that facilitates improved savings in comparison to
removing and replacing entire combustor liners or large patches or
complete panels.
[0045] Although the apparatus and methods described herein are
described in the context of cooling holes in a combustor liner of a
gas turbine engine, it is understood that the apparatus and methods
are not limited to gas turbine engines, combustor liners, or
cooling holes. Likewise, the gas turbine engine and combustor liner
components illustrated are not limited to the specific embodiments
described herein, but rather, components of both the gas turbine
engine and the combustor liner can be utilized independently and
separately from other components described herein.
[0046] While the invention has been described in terms of various
specific embodiments, those skilled in the art will recognize that
the invention can be practiced with modification within the spirit
and scope of the claims.
* * * * *