U.S. patent application number 14/269240 was filed with the patent office on 2014-08-28 for methods for repairing a turbine airfoil constructed from cmc material.
This patent application is currently assigned to General Electric Company. The applicant listed for this patent is General Electric Company. Invention is credited to Jeffrey H. Boy, Paul Edward Gray, Joseph Halada, Roger Lee Ken Matsumoto, Philip Harold Monaghan, Herbert Chidsey Roberts, III.
Application Number | 20140241900 14/269240 |
Document ID | / |
Family ID | 46551396 |
Filed Date | 2014-08-28 |
United States Patent
Application |
20140241900 |
Kind Code |
A1 |
Roberts, III; Herbert Chidsey ;
et al. |
August 28, 2014 |
METHODS FOR REPAIRING A TURBINE AIRFOIL CONSTRUCTED FROM CMC
MATERIAL
Abstract
Methods for repairing a turbine airfoil constructed from a CMC
material are provided via filling a cavity located in the turbine
airfoil with a ceramic paste (e.g., including a ceramic powder and
a binder), heating the ceramic paste in the cavity to remove the
binder, thereby forming a porous ceramic material, and adding a
molten ceramic material to the porous ceramic material. The cavity
can be defined in an airfoil of the turbine airfoil (e.g., on a tip
or cap of the airfoil). Intermediates formed during the repair of a
turbine airfoil are also provided. The intermediate can generally
include an airfoil comprising a CMC material, a cavity defined in
the airfoil, and a porous ceramic material filling the cavity.
Inventors: |
Roberts, III; Herbert Chidsey;
(Middletown, OH) ; Gray; Paul Edward; (North East,
MD) ; Matsumoto; Roger Lee Ken; (Newark, DE) ;
Boy; Jeffrey H.; (Wilmington, DE) ; Monaghan; Philip
Harold; (Hockessin, DE) ; Halada; Joseph;
(Newark, DE) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Assignee: |
General Electric Company
Schenectady
NY
|
Family ID: |
46551396 |
Appl. No.: |
14/269240 |
Filed: |
May 5, 2014 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
13188755 |
Jul 22, 2011 |
|
|
|
14269240 |
|
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|
Current U.S.
Class: |
416/241B ;
29/889.1 |
Current CPC
Class: |
C04B 41/009 20130101;
C04B 41/89 20130101; C04B 41/52 20130101; F05D 2300/6033 20130101;
F01D 5/005 20130101; F01D 5/284 20130101; C04B 41/52 20130101; C04B
41/52 20130101; C04B 41/52 20130101; C04B 41/522 20130101; Y10T
29/49318 20150115; C04B 14/4693 20130101; C04B 41/5025 20130101;
C04B 41/4582 20130101; C04B 41/4572 20130101; C04B 41/4523
20130101; C04B 41/4523 20130101; C04B 41/5059 20130101; F01D 5/282
20130101; C04B 41/4572 20130101; C04B 41/5025 20130101; C04B
41/5059 20130101; C04B 41/4539 20130101; C04B 41/4572 20130101;
C04B 41/4582 20130101; C04B 41/522 20130101; C04B 35/80 20130101;
C04B 41/4572 20130101; C04B 41/4539 20130101; C04B 41/52 20130101;
F01D 5/147 20130101; C04B 41/009 20130101 |
Class at
Publication: |
416/241.B ;
29/889.1 |
International
Class: |
F01D 5/00 20060101
F01D005/00; F01D 5/14 20060101 F01D005/14 |
Claims
1. A method of repairing an article constructed from a ceramic
matrix composite material, the method comprising: filling a cavity
located in the article with a ceramic paste, wherein the ceramic
paste comprises a ceramic powder, a ceramic fiber, and a binder;
heating the ceramic paste in the cavity to remove the binder,
thereby forming a porous ceramic material; and adding a molten
ceramic material to the porous ceramic material.
2. The method of claim 1, wherein the porous ceramic material is
heated to about 100.degree. C. or higher to remove the binder.
3. The method of claim 1, wherein the ceramic fiber comprises
silicon carbide.
4. The method of claim 3, wherein the ceramic fiber is coated with
particles.
5. The method of claim 4, wherein the particles comprise boron,
carbon, or mixtures thereof.
6. The method of claim 1, further comprising: heating the porous
ceramic material to a temperature of about 1000.degree. C. to about
1500.degree. C. prior to adding the molten ceramic material.
7. The method of claim 1, wherein the molten ceramic material is
added to the porous ceramic material at a temperature of about
1000.degree. C. to about 2000.degree. C.
8. The method of claim 1, further comprising, after adding the
molten ceramic material: cooling the porous ceramic material and
molten ceramic material to form a ceramic patch in the cavity.
9. The method of claim 8, further comprising: shaping the ceramic
patch.
10. The method of claim 1, wherein the molten ceramic material
comprises silicon carbide.
11. The method of claim 1, wherein heating the porous ceramic
material to remove the binder is achieved locally.
12. The method of claim 1, wherein the article is a turbine
airfoil.
13. A method of repairing a tip or cap of an airfoil constructed
from a ceramic matrix composite material, the method comprising:
filling a cavity located on the tip or cap of the airfoil of the
turbine airfoil with a ceramic paste, wherein the ceramic paste
comprises a ceramic powder, a ceramic fiber, and a binder; heating
the ceramic paste in the cavity to remove the binder, thereby
forming a porous ceramic material; and adding a molten ceramic
material to the porous ceramic material, wherein the molten ceramic
material comprises silicon carbide.
14. The method of claim 13, further comprising: heating the porous
ceramic material to a temperature of about 1000.degree. C. to about
1500.degree. C. prior to adding the molten ceramic material.
15. The method of claim 13, wherein the molten ceramic material is
added to the porous ceramic material at a temperature of about
1000.degree. C. to about 2000.degree. C.
16. The method of claim 13, wherein the ceramic fiber comprises
silicon carbide.
17. The method of claim 16, wherein the ceramic fiber is coated
with particles.
18. The method of claim 17, wherein the particles comprise boron,
carbon, or mixtures thereof.
19. An intermediate formed during the repair of a turbine airfoil,
the intermediate comprising: an airfoil comprising a ceramic matrix
composite material; a cavity defined in the airfoil; and a porous
ceramic material filling the cavity, wherein the porous ceramic
material comprises a ceramic powder, a ceramic fiber, and a
binder.
20. The intermediate as in claim 19, wherein the ceramic fiber
comprises silicon carbide, and wherein the ceramic fiber is coated
with particles.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] The present application claims priority to and is a
continuation application of U.S. patent application Ser. No.
13/188,755 of Herbert Chidsey Roberts III, et al. titled "Methods
for Repairing a Turbine Airfoil Constructed from CMC Material"
filed on Jul. 22, 2011, which is incorporated herein by
reference.
FIELD OF THE INVENTION
[0002] The present subject matter relates generally to repairing
tips or caps for turbine airfoils and, more particularly, to a
ceramic-based tip or cap repair for a turbine airfoil.
BACKGROUND OF THE INVENTION
[0003] In a gas turbine, air is pressurized by a compressor and
then mixed with fuel and ignited within an annular array of
combustors to generate hot gases of combustion. The hot gases flow
from each combustor through a transition piece for flow along an
annular hot gas path. Turbine stages are typically disposed along
the hot gas path such that the hot gases flow through first-stage
nozzles and airfoils and through the nozzles and airfoils of
follow-on turbine stages. The turbine airfoils may be secured to a
structural case or a plurality of rotor disks comprising the
turbine rotor, with each rotor disk being mounted to the rotor
shaft for rotation therewith.
[0004] A turbine airfoil generally includes an airfoil extending
radially outwardly from a substantially planar platform and an
attachment portion extending radially inwardly from the platform
for securing the airfoil to one of the rotor disks. Certain
airfoils can be composed of ceramic and ceramic matrix composite
(CMC) materials for operation in a high temperature environment as
exist in gas turbines. After the airfoil is put into service,
however, the tip or cap of the rotating CMC airfoil can experience
localized damage as the tip or cap comes into contact with a gas
turbine shroud or due to foreign object damage. The damage to the
CMC airfoil can lead to secondary damage if the CMC or ceramic
fibers are exposed to the moisture or other contaminates in the gas
turbine hot gas path steam.
[0005] As such, methods would be welcomed for repairing damage to a
CMC airfoil, especially on the tip or cap to extend the working
life of the airfoil.
BRIEF DESCRIPTION OF THE INVENTION
[0006] Aspects and advantages of the invention will be set forth in
part in the following description, or may be obvious from the
description, or may be learned through practice of the
invention.
[0007] Methods are generally provided for repairing an article
constructed from a CMC material. According to one exemplary method,
a cavity located in the article can be filled with a ceramic paste
(e.g., including a ceramic powder and a binder). The ceramic paste
in the cavity can then be heated to remove the binder, thereby
forming a porous ceramic material. A molten ceramic material can
then be added to the porous ceramic material. In one particular
embodiment, the cavity can be defined in an airfoil of the turbine
airfoil (e.g., on a tip or cap of the airfoil).
[0008] For example, methods are generally provided for repairing a
tip or cap of an airfoil constructed from a ceramic matrix
composite material. According to one exemplary method, a cavity
located on the tip or cap of the airfoil of the turbine airfoil can
be filled with a ceramic paste (e.g., comprising a ceramic powder
and a binder). The ceramic paste in the cavity can then be heated
to remove the binder, thereby forming a porous ceramic material. A
molten ceramic material (e.g., the molten ceramic material
comprises silicon carbide) can then be added to the porous ceramic
material.
[0009] Intermediates formed during the repair of a turbine airfoil
are also generally provided. The intermediate can generally include
an airfoil comprising a CMC material, a cavity defined in the
airfoil, and a porous ceramic material filling the cavity.
[0010] These and other features, aspects and advantages of the
present invention will become better understood with reference to
the following description and appended claims. The accompanying
drawings, which are incorporated in and constitute a part of this
specification, illustrate embodiments of the invention and,
together with the description, serve to explain the principles of
the invention.
BRIEF DESCRIPTION OF THE DRAWINGS
[0011] A full and enabling disclosure of the present invention,
including the best mode thereof, directed to one of ordinary skill
in the art, is set forth in the specification, which makes
reference to the appended figures, in which:
[0012] FIG. 1 illustrates a schematic diagram of one embodiment of
a gas turbine;
[0013] FIG. 2 illustrates a perspective view of one embodiment of a
turbine airfoil having a cavity formed in the tip or cap of the
airfoil after normal use;
[0014] FIG. 3 illustrates filling the cavity in the turbine airfoil
shown in FIG. 2 with a ceramic powder mix;
[0015] FIG. 4 illustrates an intermediate having the cavity shown
in FIG. 2 filled with a porous ceramic material after heating the
ceramic powder mix added as shown in FIG. 3;
[0016] FIG. 5 illustrates filling the porous ceramic material in
the cavity shown in FIG. 4 with a molten material; and
[0017] FIG. 6 illustrates the turbine airfoil shown in FIG. 2 after
repairing the cavity in the tip or cap.
DETAILED DESCRIPTION OF THE INVENTION
[0018] Reference now will be made in detail to embodiments of the
invention, one or more examples of which are illustrated in the
drawings. Each example is provided by way of explanation of the
invention, not limitation of the invention. In fact, it will be
apparent to those skilled in the art that various modifications and
variations can be made in the present invention without departing
from the scope or spirit of the invention. For instance, features
illustrated or described as part of one embodiment can be used with
another embodiment to yield a still further embodiment. Thus, it is
intended that the present invention covers such modifications and
variations as come within the scope of the appended claims and
their equivalents.
[0019] In general, methods are generally provided for repairing an
article (e.g., a tip or cap of an airfoil in a gas turbine), along
with the resulting repaired article (e.g., repaired airfoils). In
particular, methods are generally provided for repairing articles
(e.g., the tip or cap of an airfoil) constructed from ceramic
matrix composite (CMC) material, along with the resulting repaired
article (e.g., airfoils). Though discussed hereinafter with respect
to an airfoil, the methods may be broadly applied to any article
constructed from a CMC material.
[0020] CMC materials used to form the airfoil generally exhibit
enhanced high temperature capabilities as compared to conventional
metal-based airfoils. As such, the airfoil may reduce or eliminate
the need for supplying a cooling medium (e.g., air and the like) to
and/or through the tip or cap, thereby increasing the efficiency of
the gas turbine. Additionally, because of the low densities of CMC
materials, the weight of the tip or cap may be significantly less
than conventional metal-based tips or caps, thereby reducing the
loads generated by the tip or cap during operation of the gas
turbine. In several embodiments of the present subject mater, it
should be appreciated that the methods disclosed herein may be
designed for retrofit applications and, thus, may be configured to
be performed on pre-existing turbine airfoils.
[0021] Referring now to the drawings, FIG. 1 illustrates a
schematic diagram of a gas turbine 10. The gas turbine 10 generally
includes a compressor section 12, a combustor section 14 disposed
downstream of the compressor section 12, and a turbine section 16
disposed downstream of the combustor section 14. Additionally, the
gas turbine 10 may include a shaft 18 coupled between the
compressor section 12 and the turbine section 16. The turbine
section 16 may generally include a turbine rotor 20 having a
plurality of structural cases 22 (one of which is shown) and a
plurality of turbine airfoil 24 extending radially outwardly from
and being coupled to each structural case 22 for rotation
therewith. Each structural case 22 may, in turn, be coupled to a
portion of the shaft 18 extending through the turbine section
16.
[0022] During operation of the gas turbine 10, the compressor
section 12 pressurizes air entering the gas turbine 10 and supplies
the pressurized air to the combustors of the combustor section 14.
The pressurized air is mixed with fuel and burned within each
combustor to produce hot gases of combustion. The hot gases of
combustion flow in a hot gas path from the combustor section 14 to
the turbine section 16, wherein energy is extracted from the hot
gases by the turbine airfoils 24. The energy extracted by the
turbine airfoils 24 is used to rotate the structural cases 22 which
may, in turn, rotate the shaft 18. The mechanical rotational energy
may then be used to power the compressor section 12 and generate
electricity.
[0023] It should be understood, however, that the present
disclosure is not limited to use in structural cases 22 in the
turbine section 16 of a turbine system 10. Rather, the structural
cases 22 and/or turbine airfoils 24 may be utilized in conjunction
with any suitable section of the turbine system 10. For example,
the structural cases 22 and/or turbine airfoils 24 may, in
exemplary embodiments, be utilized in the compressor 12.
[0024] FIGS. 2-6 sequentially show an exemplary method of repairing
a tip or cap 36 of an airfoil 30 of a turbine airfoil 24.
Generally, the turbine airfoil 24 includes an attachment portion 28
and an airfoil 30 extending from a substantially planar platform
32. The platform 32 generally serves as the radially inward
boundary for the hot gases of combustion flowing through the
turbine section 16 of the gas turbine 10 (FIG. 1). The attachment
portion 28 may generally be configured to extend radially inwardly
from the platform 32 and may include a root structure (not shown),
such as a dovetail, configured to secure the airfoil 23 to the
structural case 22 of the gas turbine 10 (FIG. 1).
[0025] The airfoil 30 may generally extend radially outwardly from
the platform 32 and may include an airfoil base 34 disposed at the
platform 32 and an airfoil tip or cap 36 disposed opposite the
airfoil base 34. Thus, the airfoil tip or cap 36 may generally
define the radially outermost portion of the turbine airfoil 24.
The airfoil 30 may also include a pressure side wall 38 and a
suction side wall 40 (FIGS. 3 and 4) extending between a leading
edge 42 and a trailing edge 44. The pressure side wall 38 may
generally comprise an aerodynamic, concave outer wall of the
airfoil 30. Similarly, the suction side wall 40 may generally
define an aerodynamic, convex outer wall of the airfoil 30.
[0026] Additionally, the turbine airfoil 24 may also include an
airfoil cooling circuit 46 extending radially outwardly from the
attachment portion 28 for flowing a cooling medium, such as air,
water, steam or any other suitable fluid, throughout the airfoil
30. The airfoil cooling circuit 46 may generally have any suitable
configuration known in the art. Thus, in several embodiments, the
cooling circuit 46 may include a plurality of cooling channels or
passages extending radially within the airfoil 30, such as from the
airfoil base 34 to a location generally adjacent the airfoil tip or
cap 36. For example, in one embodiment, the airfoil cooling circuit
46 may be configured as a multiple-pass cooling circuit, with the
cooling passages being interconnected and extending radially inward
and radially outward within the airfoil 30 (e.g., in a
serpentine-like path) such that the cooling medium within the
passages flows alternately radially outwardly and radially inwardly
throughout the airfoil 30.
[0027] It should be appreciated that the various components of the
turbine airfoil 24 (e.g., the airfoil 30, platform 32 and
attachment portion 28) may generally be formed from a ceramic
matrix composite (CMC) material. In general, the CMC material used
to form the turbine airfoil 24 may comprise any suitable CMC
material known in the art and, thus, may generally include a
ceramic matrix having a suitable reinforcing material incorporated
therein to enhance the material's properties (e.g., the material
strength and/or the thermo-physical properties). In several
embodiments, the CMC material used may be configured as a
continuous fiber reinforced CMC material. For example, suitable
continuous fiber reinforced CMC materials may include, but are not
limited to, CMC materials reinforced with continuous carbon fibers,
oxide fibers, silicon carbide monofilament fibers or other CMC
materials including continuous fiber lay-ups and/or woven fiber
performs. In other embodiments, the CMC material used may be
configured as a discontinuous reinforced CMC material. For
instance, suitable discontinuous reinforced CMC materials may
include, but are not limited to, particulate, platelet, whisker,
discontinuous fiber, in situ and nano-composite reinforced CMC
materials or mixtures thereof. Moreover, it should be appreciated
that the disclosed turbine airfoil 24 may be formed from the CMC
material using any suitable manufacturing process known in the art.
For example, suitable manufacturing processes may include, but are
not limited to, injection molding, slip casting, tape casting,
infiltration methods (e.g., chemical vapor infiltration, melt
infiltration and/or the like) and various other suitable methods
and/or processes.
[0028] In normal use, defects can be formed in the turbine airfoil
24, and particularly along the airfoil 30. FIG. 2 illustrates a
perspective view of the turbine airfoil 24 having a cavity 26,
which can be the result of damage to the tip or cap. As used
herein, the term "cavity" refers to any hollow space within the
turbine airfoil 24 (e.g., in the airfoil), such as an opening,
crack, gap, aperture, hole, etc. Such a cavity 26 can be formed on
the airfoil 30 through normal use, and generally represents an area
where fragments of the original CMC material has been chipped off
of the turbine airfoil 24. As shown, the cavity 26 is located on
the tip or cap 36 of the airfoil 30, which is an area that is
particularly susceptible to such damage; however, cavities on any
portion of the turbine airfoil 24 can be repaired according to the
present disclosure.
[0029] FIG. 3 shows the addition of a ceramic paste 50 into the
cavity 26. The ceramic paste 50 can generally be applied at room
temperature (i.e., about 20.degree. C. to about 25.degree. C.) or
may be applied at an elevated temperature (e.g., up to about
100.degree. C.). In one embodiment, the ceramic paste 50 can fill
the cavity 26 completely, and may be shaped in the desired manner
to repair the cavity 26.
[0030] The ceramic paste 50 generally includes a ceramic powder, a
binder, and optionally ceramic fibers. The ceramic powder can
include silicon carbide (SiC), silicon dioxide (SiO2), Alumina
oxide (Al2O3), carbon, or mixtures thereof. The binder can include
suitable composition configured to hold the ceramic powder (and
optional ceramic fibers, if present) together as a paste, and can
include but is not limited to, an epoxy binder, a polymeric binder,
an adhesive (e.g., glue), silicon dioxide (SiO2), Alumina oxide
(Al2O3), carbon, boron, or mixtures thereof.
[0031] In one particular embodiment, the ceramic paste 50 includes
SiC and an epoxy binder. In this embodiment, SiC fibers may or may
not be included in the ceramic paste 50 if desired. The SiC fibers,
when included, can be coated to prevent absorption by the CMC
matrix exposed on the cavity 26. For example, the SiC fibers can be
coated with boron (B) or carbon (C) particles, any other suitable
particle, or mixtures thereof.
[0032] After application into the cavity 26, the ceramic paste 50
can then be heated to remove the binder from the cavity 26, leaving
a porous ceramic material 52 in the cavity 26. For example, the
ceramic paste 50 in the cavity 26 can be heated to 100.degree. C.
or greater, such as about 110.degree. C. to about 200.degree. C.
Heating can be achieved either locally (e.g., heating only the
porous ceramic material 52 and the immediate area around the cavity
26) or entirely (e.g., heating the entire airfoil 30 and/or turbine
airfoil 24). At these elevated temperatures, the binder in the
ceramic paste 50 will begin to decompose, sublimate, and/or
evaporate from the ceramic paste 50 to leave only a porous ceramic
material 52 in the cavity 26, as shown in FIG. 4. Specifically,
FIG. 4 shows an intermediary that includes a porous ceramic
material 52 (e.g., substantially free from any binder) within the
cavity 26.
[0033] The porous ceramic material 52 and cavity 26 can then be
further heated to temperatures for receiving a molten ceramic
material to fill pores in the porous ceramic material 52, thereby
forming a ceramic patch to fill the cavity 26. For example, the
porous ceramic material 52 and cavity 26 can be heated (either
locally or with the entire airfoil 30 and/or turbine airfoil 24) to
temperatures of greater than about 1000.degree. C., such as about
1100.degree. C. to about 1500.degree. C.
[0034] FIG. 5 shows molten ceramic material 54 added to the heated
porous ceramic material 52. The molten ceramic material 54 can be
added at a temperature that is greater than about 1000.degree. C.,
such as about 1100.degree. C. to about 2000.degree. C. As such, the
molten ceramic material 54 can flow and penetrate the pores of the
porous ceramic material 52, similar to in melt infusion techniques.
Additionally, the molten ceramic material 54 can bond (e.g.,
thermally, mechanically, chemically, etc.) with the porous ceramic
material 52 and/or the CMC material of the turbine airfoil 24 that
defines the cavity 26.
[0035] The molten ceramic material 54 can generally include ceramic
material, such as silicon carbide (SiC), silicon dioxide (SiO2),
Alumina oxide (Al2O3), carbon or mixtures thereof. In one
embodiment, the molten ceramic material 54 can be substantially
pure SiC (e.g., being substantially free from other compounds). As
used herein, the term "substantially free" means no more than an
insignificant trace amount present and encompasses completely free
(e.g., 0 molar % up to 0.0001 molar %).
[0036] After infusion of the molten ceramic material 54, the cavity
26 (and the rest of the turbine airfoil 24) can be cooled, allowing
the porous ceramic material 52 and molten ceramic material 54 to
cure and set into a ceramic patch 56 filling the cavity 26 (FIG.
6). As such, the cavity 26 can be filled with CMC material to
repair and refurbish the turbine airfoil. The shape of the ceramic
patch 56 can be molded as desired either during its formation at
any of the above discussed steps or after its formation upon
cooling.
[0037] As described in greater detail above, one particular
embodiment of the method of repairing a turbine airfoil constructed
from a CMC material can be summarized as follows: filling a cavity
on a CMC turbine airfoil with a ceramic paste, heating the ceramic
paste to remove the binder forming a porous structure in the
cavity, further heating the porous structure, adding molten
material to fill pores in the porous structure, and cooling the
molten material and the porous structure to form a ceramic patch in
the cavity. Other steps can be included (e.g., shaping the porous
structure, shaping the ceramic patch, etc.) in this method as
desired.
[0038] This written description uses examples to disclose the
invention, including the best mode, and also to enable any person
skilled in the art to practice the invention, including making and
using any devices or systems and performing any incorporated
methods. The patentable scope of the invention is defined by the
claims, and may include other examples that occur to those skilled
in the art. Such other examples are intended to be within the scope
of the claims if they include structural elements that do not
differ from the literal language of the claims, or if they include
equivalent structural elements with insubstantial differences from
the literal languages of the claims.
* * * * *