U.S. patent application number 13/370821 was filed with the patent office on 2012-06-07 for roughened coatings for gas turbine engine components.
This patent application is currently assigned to MT COATINGS, LLC. Invention is credited to David C. Fairbourn.
Application Number | 20120141671 13/370821 |
Document ID | / |
Family ID | 38181197 |
Filed Date | 2012-06-07 |
United States Patent
Application |
20120141671 |
Kind Code |
A1 |
Fairbourn; David C. |
June 7, 2012 |
ROUGHENED COATINGS FOR GAS TURBINE ENGINE COMPONENTS
Abstract
A gas turbine engine component with an aluminide coating on at
least a portion of an airflow surface that includes a roughening
agent effective to provide a desired surface roughness and a
deposition process for forming such aluminide coatings. A layer
including a binder and the roughening agent may be applied to the
superalloy substrate of the component and the aluminide coating
formed on the airflow surface portion by exposing the component and
layer to an appropriate deposition environment. Suitable roughening
agents include metal and ceramic particles that are dispersed on
the airflow surface portion before exposure to the deposition
environment. The particles, which are substantially intact after
the aluminide coating is formed, are dispersed in an effective
number to supply the desired surface roughness.
Inventors: |
Fairbourn; David C.; (Sandy,
UT) |
Assignee: |
MT COATINGS, LLC
Cincinnati
OH
|
Family ID: |
38181197 |
Appl. No.: |
13/370821 |
Filed: |
February 10, 2012 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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12093980 |
May 16, 2008 |
8137820 |
|
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PCT/US2006/006644 |
Feb 24, 2006 |
|
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13370821 |
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Current U.S.
Class: |
427/203 ;
427/286; 427/419.1 |
Current CPC
Class: |
Y10T 428/12535 20150115;
Y02T 50/67 20130101; Y02T 50/671 20130101; Y02T 50/673 20130101;
Y10T 428/12993 20150115; F05D 2250/61 20130101; F05D 2300/611
20130101; Y10T 428/12736 20150115; F01D 5/288 20130101; F05D
2300/21 20130101; F05D 2230/31 20130101; F05D 2250/28 20130101;
F05D 2230/90 20130101; F05D 2300/614 20130101; Y02T 50/60
20130101 |
Class at
Publication: |
427/203 ;
427/419.1; 427/286 |
International
Class: |
B05D 7/14 20060101
B05D007/14; B05D 5/00 20060101 B05D005/00; B05D 1/36 20060101
B05D001/36 |
Claims
1. A deposition process for a superalloy gas turbine engine
component having an airflow surface, comprising: dispersing a
plurality of particles on at least a portion of the airflow
surface; and forming an aluminide coating on the airflow surface
portion that includes the dispersed particles in a substantially
intact condition and in an effective number such that the aluminide
coating has a desired surface roughness.
2. The deposition process of claim 1 wherein forming the aluminide
coating further comprises: exposing the gas turbine engine
component to a deposition environment effective to form the
aluminide coating.
3. The deposition process of claim 2 wherein dispersing the
particles further comprises: applying a layer comprising the
particles on the airflow surface portion and a binder effective to
adhere the particles to the airflow surface while the aluminide
coating is formed.
4. The deposition process of claim 3 wherein exposing the gas
turbine engine component to the deposition environment further
comprises: forming the aluminide coating from a metal originating
from a vaporized donor material, silicon from the layer, and the
particles from the layer.
5. The deposition process of claim 1 wherein dispersing the
particles further comprises: applying a layer comprising the
particles on the airflow surface portion and a binder effective to
adhere the particles to the airflow surface while the aluminide
coating is formed.
6. The deposition process of claim 5 wherein applying the layer
further comprises: applying the layer across an entire surface area
of the airflow surface.
7. The deposition process of claim 5 wherein applying the layer
further comprises: applying the layer in a discrete areas on the
airflow surface.
8. The deposition process of claim 5 wherein the substrate includes
first and second edges bounding the airflow surface, and applying
the layer further comprises: applying stripes of the layer that are
inclined to intersect at least one of the first and second
edges.
9. The deposition process of claim 5 wherein the substrate includes
first and second edges bounding the airflow surface, and applying
the layer further comprises: applying stripes of the layer that are
substantially parallel to the first and second edges.
10. The deposition process of claim 5 wherein applying the layer
further comprises: mixing the particles with a binder comprising a
silane solution; and placing the binder on the airflow surface
portion so as to bind the particles to the airflow surface
portion.
11. The deposition process of claim 1 wherein the desired surface
roughness of the aluminide coating is greater than about 0.68
microinches.
12. The deposition process of claim 1 wherein the desired surface
roughness of the aluminide coating is greater than about 0.75
microinches.
13. The deposition process of claim 1 wherein the desired surface
roughness of the aluminide coating is greater than about 100
microinches.
14. The deposition process of claim 1 wherein the desired surface
roughness of the aluminide coating ranges from about 120
microinches to about 130 microinches.
15. The deposition process of claim 1 wherein the airflow surface
has a convex curvature.
16. The deposition process of claim 1 wherein the particles are
distributed in separated discrete areas across the airflow surface.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application is a divisional of U.S. application Ser.
No. 12/093,980, filed May 16, 2008, which is the National Stage of
International Application No. PCT/US2006/006644, filed Feb. 24,
2006. The content of each of these applications is hereby
incorporated by reference herein in its entirety for all
purposes.
FIELD OF THE INVENTION
[0002] The present invention relates to coated metal components
and, more particularly, gas turbine engine components with a
roughened coating and methods of forming such roughened coatings on
gas turbine engine components.
BACKGROUND OF THE INVENTION
[0003] Intermetallic layers and coatings are often formed on a
surface of a metal component to protect the underlying metal
substrate of the component and to extend its useful life during
operation. For example, many superalloy components in gas turbine
engines, like turbine blades, vanes, and nozzle guides, include an
aluminide coating on airflow surfaces that protects the underlying
superalloy base metal from high temperature oxidation and
corrosion. Among other applications, gas turbine engines are used
as aircraft or jet engines, such as turbofans. Gas turbine engines
are also used in electromotive power generation equipment, such as
industrial gas turbine engines, to generate electricity, and as
power plants providing motive forces to propel vehicles.
[0004] Generally, gas turbine engines include a compressor for
compressing air, a combustor for mixing the compressed air with
fuel, such as jet fuel or natural gas, and igniting the mixture,
and a turbine blade assembly for producing power. In particular,
gas turbine engines operate by drawing air into the front of the
engine. The air is then compressed, mixed with fuel, and combusted.
Hot exhaust gases from the combusted mixture pass through a
turbine, which causes the turbine to spin about an axial center and
thereby powers the compressor. Aircraft gas turbine engines,
referred to herein as jet engines, propel the attached aircraft in
response to the thrust provided by the flow of the hot exhaust
gases from the gas turbine engine. Rotation of the turbine in
industrial gas turbine engines generates electrical power and
motive power for vehicles.
[0005] Gas turbine engines include turbine blades shaped as
airfoils and coupled to the turbine. The hot exhaust gases from the
combustor flow over and under each turbine blade. Because of the
airfoil shape, the flow path across the top of the airfoil or
convex side is much longer than the flow path underneath the
concave side of the turbine blade. The result is an aerodynamic
lift, which drives each of the turbine blades in the desired
direction. Work is then extracted from the coordinated rotation of
the turbine blades about the axial center of the gas turbine.
[0006] Conventional approaches for optimizing aerodynamic lift
generated by the spinning turbine blades rely on increasingly
radical airfoil shapes and three-dimensional topologies. However,
these conventional approaches that focus solely upon advances in
component geometry introduce complexity into the component
manufacture process and are ultimately limited in the improvement
in aerodynamic efficiency.
[0007] Accordingly, there is a need for gas turbine engine
components with improved lift and methods of forming such gas
turbine engine components that avoids the necessity of a complex
airfoil shape.
SUMMARY OF INVENTION
[0008] The present invention provides, in one aspect, an airflow
surface of a gas turbine engine component is at least partially
covered with an aluminide coating including an effective number of
substantially-intact particles dispersed therein such that the
aluminide coating has a desired or targeted surface roughness. The
gas turbine engine component is formed from a superalloy material,
such as a nickel-based superalloy. The gas turbine engine component
may be a turbine blade for a gas turbine engine and, in particular,
a jet engine turbine blade for a jet turbine engine.
[0009] Advantageously, the aluminide coating on the airflow surface
portion may be formed by a deposition process that includes
dispersing the particles on at least the portion of the airflow
surface and then forming the aluminide coating that includes the
dispersed particles in a substantially intact condition and in an
effective number such that the aluminide coating has a desired or
targeted surface roughness. The method may include applying a layer
containing silicon and the particles, such as a mixture of silane
and either ceramic or metallic particles, to at least the portion
of the airflow surface. After applying the layer, the gas turbine
engine component is exposed to a deposition environment effective
to form the aluminide coating with the dispersed particles. One
suitable deposition environment relies on vaporizing a donor
material including a metal effective to form the aluminide layer,
which includes the metal from the donor material, silicon from the
layer, and the particles from the layer.
[0010] The surface finish of the present invention deviates from
conventional turbine blade designs that want the surface finishes
on the entire airflow surface to be substantially identical. In
contrast, the present invention provides a surface finish on one
portion of the airflow surface (i.e., the convex airflow surface
found on most gas turbine blades) that differs from the surface
finish on another portion of the air flow surface (e.g., the
opposite concave airflow surface found on most gas turbine
blades).
[0011] The surface finish of the present invention deviates from
conventional turbine blade designs that specify the turbine blades
to be as smooth as possible to contribute to laminar flow and to
optimize the flow of hot exhaust gases beneath the concave portion
of the airflow surface of the turbine blade. Typically, a desired
surface roughness (R.sub.A) for the surface finishes of the convex
and concave portions of the airflow surface is less than about 68
microinches, after aluminiding. In contrast, the present invention
advantageously applies an aluminide coating to the convex airflow
surface portion that increases the surface roughness above this
conventional desired value. The concave airflow surface portion may
have a conventional surface roughness but, in any event, has a
smoother surface than all or part of the convex airflow surface
portion. The difference in surface roughness slows the airflow
velocity across the convex airflow surface portion in comparison to
the airflow velocity across the concave airflow surface
portion.
[0012] The present invention improves the aerodynamic efficiency of
gas turbine engine components providing aerodynamic lift without
the need for complex component geometries and/or improves the
aerodynamic lift in components having complex geometries beyond the
gains provided solely by the geometry.
[0013] These and other benefits and advantages of the present
invention shall be made apparent from the accompanying drawings and
description thereof.
BRIEF DESCRIPTION OF THE DRAWINGS
[0014] The accompanying drawings, which are incorporated in and
constitute a part of this specification, illustrate an embodiment
of the invention and, together with a general description of the
invention given above, and the detailed description of the
embodiment given below, serve to explain the principles of the
invention.
[0015] FIG. 1 is a perspective view of a gas turbine engine
component with a liquid being applied to a portion of the gas
turbine engine component in accordance with the principles of the
present invention;
[0016] FIG. 1A is a diagrammatic cross-sectional view of a portion
of the coated gas turbine engine component of FIG. 1;
[0017] FIG. 2 is a schematic view showing gas turbine engine
components, such as that from FIG. 1, in a deposition environment
of a simple CVD deposition system for purposes of explaining the
principles of the present invention;
[0018] FIGS. 3A-C are perspective views similar to FIG. 1 in
accordance with alternative embodiments of the invention;
[0019] FIG. 4 is a diagrammatic cross-sectional view of a portion
of a coated gas turbine engine component of the present invention;
and
[0020] FIG. 4A is a diagrammatic cross-sectional view of a portion
of a coated gas turbine engine component in accordance with an
alternative embodiment of the present invention.
DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS
[0021] With reference to FIG. 1 and in accordance with the
principles of the present invention, a gas turbine engine component
10, in a representative construction, includes an airfoil segment
12 designed to be in the high-pressure, hot airflow path (as
indicated by arrows 13). Integral with airfoil segment 12 is a root
14 used to secure gas turbine engine component 10 to the turbine
disk (not shown) of the gas turbine engine (not shown). The airfoil
segment 12 is fabricated from any suitable nickel-, cobalt-, or
iron-based high temperature superalloy from which such gas turbine
engine components 10 are commonly made. The base element, typically
nickel or cobalt, is by weight the single greatest element in the
superalloy. For example, where the component 10 is a gas turbine
component in a jet engine, segment 12 may be the nickel-based
superalloy Inconel 795 Mod5A or CMSX-4. The present invention is,
however, not intended to be limited to any particular gas turbine
engine component 10, and may be any high pressure turbine blade,
low pressure turbine blade, or any other component of a gas turbine
having an airfoil surface that generates lift while operating in a
jet engine or while operating in an industrial gas turbine
engine.
[0022] A surface 16 of the airfoil segment 12 of gas turbine engine
component 10 is divided into airflow surfaces 18, 20 extending
between a curved tip edge 22 and a curved foil tip edge 24. Cooling
channels or passages internal to airfoil segment 12 include surface
cooling holes 26 on surface 16 so as to permit cooling air to pass
through the interior of airfoil segment 12 while gas turbine engine
component 10 is in service on the gas turbine engine. The root 14
includes a contoured surface 28 extending beneath a platform 30 and
is separated from the airfoil segment 12 by the platform 30.
[0023] Depending upon the use of the gas turbine engine component
10, combustion gases in the airflow path 13 may have a temperature
as high as 3000.degree. F. This promotes heating of the airfoil
segment 12. Gas cooling of the airfoil segment 12 limits operating
temperatures to 1800.degree. F. or less. When the gas turbine
engine component 10 is in service, portions of the component 10
below the platform 30 are cooler than the airfoil segment 12 and,
frequently, are at an operating temperature of less than
1500.degree. F. when the component 10 is in service. The cooler
portions include the root 14, which is coupled with an air-cooled
turbine disk of the gas turbine.
[0024] Airflow surface 20 has a concave shape extending between
leading edge 22 and trailing edge 24 and airflow surface 18 has a
convex shape extending between edges 22 and 24. When the gas
turbine engine component 10 is coupled with a gas turbine engine
(not shown) and rotated, leading edge 22 is the first to encounter
the hot exhaust gases and air in the airflow path 13. The airflow
path 13 will split at edge 22. A portion of the hot exhaust gases
and air in air flow path 13 will flow across airflow surface 18 and
another portion of the hot exhaust gases and air will flow across
airflow surface 20. Due to the difference in curvature and length,
the flow velocity will be greater across airflow surface 18 than
across airflow surface 20. Due to the familiar Bernoulli's
principle, lift is generated because the pressure is greater near
airflow surface 20 than near airflow surface 18. The split airflow
recombines after passing trailing edge 24.
[0025] With reference to FIGS. 1 and 1A, a layer 40 is applied to
all or a portion of convex airflow surface 18 of gas turbine engine
component 10 before an aluminide coating 42 (FIG. 4) is formed on
selected regions of convex airflow surface 18 in a CVD apparatus 50
(FIG. 2). The layer 40 may be applied as a liquid or solution and
then dried to remove solvent and form a solid or semi-solid coating
on the gas turbine engine component 10 before aluminiding
occurs.
[0026] The layer 40 is applied to all or a portion of convex
airflow surface 18, such as by hand application with a paintbrush B
or another type of conventional applicator recognized by a person
having ordinary skill in the art. Alternatively, gas turbine engine
component 10 may be sprayed with a suitable liquid or solution
before drying and aluminiding. Thereafter, the coated gas turbine
engine component 10 (which may advantageously first be dried and
heated) is placed into a deposition environment 52 (FIG. 2)
whereupon the aluminide coating 42 will be formed on convex airflow
surface 18 to the desired thickness.
[0027] The layer 40 applied to all or a portion of the convex
airflow surface 18 is initially a liquid or solution that includes
a binder 38 and a roughening agent, such as inorganic particles 44,
blended with the binder 38. The liquid forming the binder 38 may be
a silicon-containing binder such as a silane and, advantageously,
may be a high-viscosity silane. Silanes suitable for use in the
present invention may have mono-, bis-, or tri-functional trialkoxy
silane. The silane may be a bifunctional trialkoxy silyl,
preferably trimethoxy, or triethoxy silyl groups. Amino silanes may
also be used, although thio silanes may not be desired due to their
sulfur content. Bisfunctional silane compounds are well known, and
two suitable for use in the present invention are
bis(triethoxysilyl) ethane and bis(trimethoxysilyl) methane. In
both of these compounds, the bridging group between the two silane
moieties is an alkyl group. Additional commercially available
silanes include, but are not limited to, [0028]
1,2-Bis(tetramethyldisoloxanyl) Ethane [0029]
1,9-Bis(triethoxysilyl) Nonane [0030] Bis(triethoxysilyl) Octane
[0031] Bis(trimethoxysilyl Ethane [0032]
1,3-Bis(trimethylsiloxy)-1,3-Dimethyl Disiloxane [0033]
Bis(trimethylsiloxy) Ethylsilane [0034] Bis(trimethylsiloxy)
Methylsilane [0035] A1-501 available from AG Chemetall (Frankfurt
Germany)
[0036] The silane of binder 38 may be neat, in an aqueous solution,
or in an aqueous/alcohol solvent solution. A solvent for the latter
type of solution may contain from about 1% to 2% by volume (vol. %)
to about 30 vol. % deionized water with the remainder of the
solution being a lower alcohol, such as methanol, ethanol,
isopropanol, or the like. The solvent is combined with the silane
and glacial acetic acid to establish a pH of about 4 to 6. The
concentration of the silane compound is not relevant as long as the
silane remains in solution during application. Generally, the
solution will include about 1% to about 20% silane, which may be
measured either by volume or by weight in this concentration
range.
[0037] The binder 38 of layer 40 applied to gas turbine engine
component 10 is allowed to dry and then is heated, such as with a
heat gun (not shown) or in a heated enclosure (not shown), to a
temperature suitable to release or remove solvent from the binder
38 and provide a solid or semisolid cured state. Before curing, the
layer 40 on the convex airflow surface 18 may first be allowed to
dry, such as underneath a lamp (not shown), to partially remove the
constituent solvent. Generally, the layer 40 is applied in an
amount of about 0.01 g/cm.sup.2 to about 2.0 g/cm.sup.2. Multiple
layers 40 of liquid or solution may be applied to convex airflow
surface 18, each individual layer 40 being dried and heated before
applying the next successive layer 40. As used herein, the layer 40
may refer to either the initially applied layer of liquid or
solution or, without limitation, to the cured or dried layer that
has had solvent removed from binder 38 by heating and/or air curing
at room temperature.
[0038] The particles 44 of layer 40 constituting the roughening
agent may advantageously be composed of a ceramic, such as silica,
alumina, chromium dioxide, yttria, hafnia, zirconia, and
combinations and mixtures thereof. For example, the particles 44
may be a fine alumina flour having a mesh size on the order of 270
to 325 mesh or finer. Alternatively, the particles 44 may include a
metal, such as boron, aluminum, chromium, yttrium, hafnium,
zirconium, and combinations and alloys thereof. Alternatively, the
particles 44 may be a metallic powder comprised of metallurgy
identical to the base metal constituting a substrate 46 of the gas
turbine engine component 10 and with an optional addition of less
than about 1% by weight of boron powder. Preferably, the layer 40
is not allowed to infiltrate into the cooling holes 26 during
application to the gas turbine engine component 10. The binder 38
of layer 40, after curing, secures the particles 44 to the airflow
surface 18 during the aluminiding process. The invention
contemplates other types or compositions of binders 38, which may
lack a silicon content, may be used to retain the dispersed
particles 44 on airflow surface 18 before aluminiding.
[0039] With reference to FIG. 2, a CVD apparatus 50 suitable for
use in forming the aluminide coating 42 (FIG. 4) includes a main
reaction chamber 54 enclosing the interior space defining a
deposition environment 52 when purged of atmospheric gases, and
evacuated. Inert gas, such as argon, is supplied from a gas supply
56 to the reaction chamber 54 through an inlet port 58 defined in
the wall of chamber 54. An exhaust port 60 defined in the wall of
the reaction chamber 54 is coupled with a vacuum pump 62 capable of
evacuating the reaction chamber 54 to a vacuum pressure. One or
more gas turbine engine components 10 are introduced into the
reaction chamber 54 and are situated away from a source of
extrinsic metal, as explained below.
[0040] Positioned within the reaction chamber 54 is a mass or
charge of a solid donor material 64, a mass or charge of an
activator material 66, and several gas turbine engine components
10. Suitable solid donor materials 64 include alloys of chromium
and aluminum, which are preferably low in sulfur content (<3 ppm
sulfur). One suitable donor material 64 is 44 wt % aluminum and
balance chromium. Appropriate activator materials 66 suitable for
use in the invention include, but are not limited to, aluminum
fluoride, aluminum chloride, ammonium fluoride, ammonium chloride,
and ammonium bifluoride. The reaction chamber 54 is heated to a
temperature effective to cause vaporization of the activator
material 66, which is transported as diagrammatically indicated by
arrows 65 within the deposition environment 52 to the solid donor
material 64. Typically, this temperature ranges from about
1950.degree. F. to about 2000.degree. F. Interaction between the
vaporized activator material 66 and the solid donor material 64
promotes the release of a vapor phase reactant from the solid donor
material 64. This vapor contains an extrinsic metal, typically
aluminum, that contributes a first extrinsic metal for
incorporation into an aluminide coating 42 (FIG. 4) formed on
component 10, as diagrammatically indicated by arrows 68. The
extrinsic metal is separate, distinct, and independent from the
material comprising the gas turbine engine component 10 and any
coating preapplied to component 10.
[0041] With reference to FIGS. 3A-C in which like reference
numerals refer to like features in FIG. 2 and in accordance with
alternative embodiments of the invention, layer 40 may be provided
on only selected regions of the convex airflow surface 18. As shown
in FIG. 3A, layer 40 may be applied in discrete areas distributed
across the convex airflow surface 18 as a plurality of
substantially-parallel stripes 70 having either smooth edges or
jagged, uneven edges. The stripes 70 of layer 40 are aligned
substantially parallel to the leading and trailing edges 22, 24
that bound the airflow surface 18.
[0042] As shown in FIG. 3B, layer 40 may be applied across the
convex airflow surface 18 in discrete areas defined by a plurality
of discrete islands or areas 72 arranged either randomly or in
specific rows and/or columns. The peripheral boundary surrounding
the discrete areas 72 of layer 40 may be irregular, as shown,
angular, curvilinear, regular (e.g., circular), or a distribution
of different shapes.
[0043] As shown in FIG. 3C, the layer 40 may be applied across the
convex airflow surface 18 as a pattern of substantially-parallel
stripes 74 inclined diagonally across the convex airflow surface
18. The resultant airflow path is believed to occur in channels
defined between adjacent stripes 74 and is directed toward the root
14 of the component 10, which represents a non-airflow surface.
Each of the stripes 74 of layer 40 intersects at least one of the
first and second edges 22, 24 that bound the airflow surface 18.
This particular pattern for layer 40 may cause the air in airflow
path 13 to twist as it tumbles through the gas turbine engine (not
shown).
[0044] With reference to FIG. 4, the aluminide coating 42 is formed
on a metallic substrate 46 of the gas turbine engine component 10
across at least the airflow surfaces 18, 20. The aluminide coating
42 on the convex airflow surface 18 on areas with the pre-applied
layer 40 will include the particles 44 and may include one or more
elements from the binder 38 (FIG. 1A). The spatial distribution of
the particles 44 determines the topography of the aluminide coating
42 in these areas. Advantageously, the viscosity of the binder 38
is sufficient to cover the particles 44 so that the particles 44
are buried or submerged in the binder 38 before aluminiding.
[0045] The particles 44 operate to effectively increase the surface
roughness of the aluminide coating 42 in comparison with adjacent
portions of convex airflow surface 18, if any, lacking layer 40
before aluminiding. Particles 44 create raised or elevated surface
irregularities or mounds in the aluminide coating 42 at distributed
locations across the convex airflow surface 18. This difference in
surface finish is best apparent from FIG. 4 as areas of the
aluminide coating 42 proximate or local to each particle 44 will
have an average thickness of h.sub.1, as compared to the nominal
thickness, h.sub.0, of the aluminide coating 42 in regions between
adjacent particles 44 and not affected by the presence of the
particles 44. Of course, the increase in average surface roughness
will also reflect the number or density of particles 44 and will
statistically include portions of aluminide coating 42 overlying
the particles 44 and having local effective thicknesses ranging
between h.sub.1 and h.sub.0 due to the mounding. The thickness
h.sub.0 may be thicker than the thickness of the aluminide coating
42 formed on other surfaces of the gas turbine engine component 10,
such as on concave airflow surface 20, because of the presence of
silicon originating from binder 38.
[0046] The particles 44 remain substantially intact after the
aluminiding process forming the aluminide coating 42. Preferably,
the particles 44 originally dispersed in the pre-applied layer 40
are incorporated into the aluminide coating 42 without significant
degradation by the aluminiding process or at the temperature of the
aluminiding process. The number of particles 44 dispersed in the
aluminide coating 42 is effective to provide the aluminide coating
42 with a desired surface roughness. The value of the average or
peak surface roughness is contingent upon, among other parameters,
the size, shape, distribution, and number of particles 44 dispersed
in the aluminide coating 42. Preferably, the surface finish of
aluminide coating 42 has an average surface roughness (R.sub.A)
greater than a conventional surface finish, considered to lack
particles similar to particles 44, of about 68 microinches.
Advantageously, the average surface roughness of aluminide coating
42 is greater than about 75 microinches. More advantageously, the
average surface roughness of aluminide coating 42 is greater than
about 100 microinches. Most advantageously, the particles 44
influence the aluminide coating 42 to provide an average surface
roughness that ranges from about 120 microinches to about 130
microinches.
[0047] The particles 44 are illustrated in FIG. 4 as having a
substantially uniform size. However, the invention is not so
limited as particles 44 may have a distribution of sizes with a
size range effective provide the desired surface finish on airflow
surface 20. The particles 44 are illustrated in FIG. 4 as being
approximately spherical. However, the invention is not so limited
as particles 44 may have other appropriate three-dimensional
geometrical shapes, such as elongated cylinders, rods, needles,
pyramids, etc. The particles 44 are illustrated in FIG. 4 as being
positioned approximately at the position of the original concave
airflow surface 18. However, the invention is not so limited as
particles 44 may be positioned with a distribution of locations
across the thickness of aluminide coating 42 between the surfaces
18 and 45.
[0048] In this specific embodiment of the present invention,
aluminide coating 42 operates as an environmental coating having a
working surface 45 exposed to the atmosphere with the gas turbine
engine component 10 in service. The general composition of
aluminide coating 42 in regions of the convex airflow surface 18
initially covered by layer 40 may advantageously include a
concentration of silicon if the binder 38 contains silicon. In this
instance, the concentration of silicon in the aluminide coating 42
may be, for example, about 0.5 percent by weight (wt %).
[0049] The presence of silicon in the aluminide coating 42 may also
increase the thickness of the aluminide coating 42 in regions of
the convex airflow surface 18 initially covered by layer 40, in
comparison with the aluminide coating 42 on regions of the convex
airflow surface 18 not initially covered by layer 40. This
increased comparative thickness may also effectively contribute to
the roughening of the convex airflow surface 18 if the layer 40 is
applied to selected regions, as shown for example in FIGS.
3A-C.
[0050] With reference to FIG. 4A, the invention contemplates
aluminide coating 42 may partially diffuse into the substrate 46
beneath the original convex airflow surface 18 of the substrate 46,
instead of being a purely additive layer as shown in FIG. 4. The
resulting aluminide coating 42 includes a diffusion region 41 that
extends beneath the formed position of the original convex airflow
surface 18 and an additive region 43 overlying the former position
of the original convex airflow surface 18. In this instance, the
outermost boundary of the additive region 43 defines the working
surface 45 of aluminide coating 42 when the gas turbine engine
component 10 is in service. Additive region 43 is an alloy that
includes a relatively high concentration of the donor metal
aluminum and a concentration of a metal, for example nickel, from
substrate 46 outwardly diffusing from component 10. By contrast,
diffusion region 41 has a lower concentration of aluminum and a
relatively high concentration of the metal of substrate 46.
[0051] The present invention may be used in combination with the
application of a platinum aluminide coating on gas turbine engine
component 10. In this instance, layer 40 is placed on the gas
turbine engine component 10 after the coating of platinum but
before aluminiding.
[0052] The aluminide layer 42 containing particles 44 may also be
formed on gas turbine engine components 10 including the
silicon-containing layer 30 by various alternative techniques known
in the art, including but not limited to dynamic CVD and pack
coating deposition processes such as an above-the-pack process or
an in-the-pack process or by electrospark deposition or
alloying.
[0053] The present invention is generally applicable to turbine
engine components 10 used in the gas turbines of jet engines, the
gas turbines of industrial gas turbine engines, or in other
turbomachinery. In particular, the present invention is applicable
for roughening turbine blades in such engines and, more
particularly, for roughening turbine blades in the gas turbines
used in jet engines.
[0054] While the present invention has been illustrated by the
description of an embodiment thereof and specific examples, and
while the embodiment has been described in considerable detail, it
is not intended to restrict or in any way limit the scope of the
appended claims to such detail. Additional advantages and
modifications will readily appear to those skilled in the art. The
invention in its broader aspects is therefore not limited to the
specific details, representative apparatus and methods and
illustrative examples shown and described. Accordingly, departures
may be made from such details without departing from the scope or
spirit of applicant's general inventive concept.
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