U.S. patent application number 15/686803 was filed with the patent office on 2018-03-01 for hot corrosion-resistant coatings for gas turbine components.
The applicant listed for this patent is BARSON COMPOSITES CORPORATION. Invention is credited to Charles Clifford BERGER, David John WORTMAN.
Application Number | 20180058228 15/686803 |
Document ID | / |
Family ID | 61241883 |
Filed Date | 2018-03-01 |
United States Patent
Application |
20180058228 |
Kind Code |
A1 |
BERGER; Charles Clifford ;
et al. |
March 1, 2018 |
HOT CORROSION-RESISTANT COATINGS FOR GAS TURBINE COMPONENTS
Abstract
A gas turbine component for use in a gas turbine engine includes
a substrate a ceramic-based thermal barrier coating (TBC), and a
diffusion chromide bond coat between the base material and the TBC.
A thermally grown oxide (TGO) layer can be formed on the bond coat
prior to application of the TBC. The TBC and the TGO include a
common metal oxide. The oxide can be sacrificially in use and
soluble in a molten sulfate salt, make the coating system
particularly suitable for use in a marine environment.
Inventors: |
BERGER; Charles Clifford;
(New City, NY) ; WORTMAN; David John; (Hamilton,
OH) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
BARSON COMPOSITES CORPORATION |
Old Bethpage |
NY |
US |
|
|
Family ID: |
61241883 |
Appl. No.: |
15/686803 |
Filed: |
August 25, 2017 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
62380029 |
Aug 26, 2016 |
|
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|
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
C23C 10/10 20130101;
Y02T 50/60 20130101; F05D 2300/134 20130101; F05D 2300/182
20130101; F05D 2300/5023 20130101; C23C 10/60 20130101; F04D 29/324
20130101; F05D 2300/143 20130101; F05D 2220/32 20130101; F05D
2240/20 20130101; F05D 2300/135 20130101; C23C 28/30 20130101; C23C
28/32 20130101; C23C 28/3455 20130101; F05D 2230/90 20130101; F01D
11/00 20130101; F05D 2300/132 20130101; C23C 28/322 20130101; F04D
29/542 20130101; F05D 2300/175 20130101; Y02T 50/6765 20180501;
F01D 5/288 20130101; C23C 10/14 20130101; F01D 9/02 20130101; F05D
2240/60 20130101; F05D 2300/222 20130101; F05D 2230/31 20130101;
F05D 2240/12 20130101 |
International
Class: |
F01D 5/28 20060101
F01D005/28; C23C 10/10 20060101 C23C010/10; C23C 10/60 20060101
C23C010/60 |
Claims
1. A gas turbine component for use in a gas turbine engine,
comprising: a substrate comprising a metal base material; a
ceramic-based thermal barrier coating disposed over the substrate,
the thermal barrier coating defining at least a portion of an outer
surface of the gas turbine component; and a bond coat disposed
between the base material and the thermal barrier coating, the bond
coat comprising a chromide diffusion coating.
2. A gas turbine component as defined in claim 1, wherein the bond
coat comprises a platinum diffusion coating.
3. A gas turbine component as defined in claim 1, wherein the bond
coat comprises an aluminide diffusion coating.
4. A gas turbine component as defined in claim 1, wherein the bond
coat further comprises hafnium, silicon, zirconium, or any
combination thereof.
5. A gas turbine component as defined in claim 1, wherein the bond
coat is a diffusion coating comprising chromium, platinum,
aluminum, and at least one of hafnium, silicon or zirconium.
6. A gas turbine component as defined in claim 5, wherein the base
material is a Ni-based superalloy having a gamma phase and a gamma
prime phase distributed within the gamma phase, the platinum of the
diffusion coating residing in the gamma prime phase and the
chromium of the diffusion coating residing in the gamma phase.
7. A gas turbine component as defined in claim 1, wherein the
component is a gas turbine blade comprising an airfoil that is
exposed to combustion gases of a gas turbine engine when in use,
the bond coat and the thermal barrier coating being located along
the airfoil.
8. A method of making a gas turbine component having a thermal
barrier coating defining at least a portion of an outer surface of
the gas turbine component, the method comprising the step of
forming a diffusion bond coat on a component substrate, the
diffusion bond coat comprising chromium interdiffused with a metal
substrate material of the component substrate, wherein the thermal
barrier coating is subsequently coated over the bond coat.
9. The method of claim 8, wherein the diffusion bond coat comprises
a Pt-aluminide coating.
10. The method of claim 8, wherein the step of forming the
diffusion bond coat comprises vapor phase deposition of the
chromium on the metal substrate material.
11. The method of claim 8, wherein the diffusion bond coat
comprises hafnium, silicon, zirconium, or any combination
thereof.
12. The method of claim 8, wherein the step of forming the
diffusion bond coat comprises the steps of coating a slurry
comprising a platinum-group metal over the substrate material and
heat treating the slurry-coated substrate material to interdiffuse
the platinum-group metal with the substrate material.
13. The method of claim 12, wherein the slurry further comprises
hafnium, silicon, zirconium, or any combination thereof.
14. The method of claim 12, wherein the step of forming the
diffusion bond coat further comprises vapor phase deposition of the
chromium on the metal substrate material before the step of coating
the slurry over the substrate material.
15. The method of claim 12, wherein the slurry comprises the
chromium of the bond coat, the chromium being interdiffused with
the substrate material during the step of heat treating.
16. The method of claim 12, further comprising the step of vapor
phase aluminide coating the component substrate after the step of
heat treating, whereby the diffusion bond coat further comprises an
aluminide coating.
17. The method of claim 8, further comprising the step of forming a
thermally grown oxide layer over the diffusion bond coat, wherein
the thermal barrier coating is subsequently coated over the
thermally grown oxide layer.
18. A gas turbine component for use in a gas turbine engine,
comprising: a substrate comprising a metal base material; a metal
bond coat formed on the base material; a thermally grown oxide
layer formed on the bond coat and comprising an oxide of a metal
element of the bond coat; and a ceramic-based thermal barrier
coating disposed over the substrate and defining at least a portion
of an outer surface of the gas turbine component, wherein the
thermal barrier coating further comprises the oxide of the metal
element of the bond coat.
19. A gas turbine component as defined in claim 18, wherein the
oxide of the metal element is alumina or chromia.
20. A gas turbine component as defined in claim 18, wherein the
bond coat comprises a diffusion coating comprising chromide,
Pt-aluminide, and at least one of hafnium, silicon, or
zirconium.
21. A method of making a gas turbine component comprising the step
of coating a turbine component substrate with a coating system
comprising a ceramic-based thermal barrier coating comprising a
sacrificial oxide that is soluble in a molten sulfate salt.
Description
TECHNICAL FIELD
[0001] This disclosure generally relates to coatings and surface
treatments for gas turbine components.
BACKGROUND
[0002] Certain critical gas turbine components, particularly those
in the hot section of a gas turbine, are exposed to a harsh
operating environment that may expose the component to high
temperatures, high mechanical stresses, and potentially reactive
combustion gases. The possible effects of this type of operating
environment may be considered when selecting turbine component
materials. For example, material characteristics such as resistance
to heat, stress, fatigue, corrosion, erosion, and/or oxidation may
be considered. Material costs and manufacturability may be
considered as well, along with numerous other factors. While many
advancements have been made in turbine component materials and
coatings in the aviation industry, they have failed to address
certain aspects of the operating environment of non-aviation
applications, such as marine vessel powerplant applications. In
such applications, airborne salts contribute to corrosion, and peak
operating temperatures may be sustained for longer periods than in
applications in which the turbine is primarily used for propulsion.
Exposure of the turbine component to turbine fuels, some of which
include contaminants such as vanadium or sulfur, can also
contribute to corrosion.
[0003] These types of conditions will likely cause turbine
components in future naval gas turbine engines to have reduced
service life if not addressed, for example from about 20,000 hours
currently to about 10,000 hours or less, due to elevated
temperatures and the effects of sodium and calcium sulfate
deposition on turbine component surfaces. This could lead to
shorter intervals for engine removal and have a significant impact
on the readiness of the U.S. Navy surface fleet while also
significantly increasing costs (e.g., maintenance, repair,
replacement, etc.). These next generation gas turbines may use
thermal barrier coatings (TBCs) on turbine blades and vanes to
improve engine fuel efficiency and increase horsepower. In
addition, the engines may need to operate at higher power levels to
support electrical generation for other shipboard uses. The effect
will be an increased percentage of time at peak temperatures
compared to current levels and/or increased turbine component
temperatures even with the use of TBCs to reduce component
substrate temperatures.
SUMMARY
[0004] In accordance with various embodiments, a gas turbine
component for use in a gas turbine engine includes a substrate
comprising a metal base material, a ceramic-based thermal barrier
coating disposed over the substrate, and a bond coat disposed
between the base material and the thermal barrier coating. The
thermal barrier coating defines at least a portion of an outer
surface of the gas turbine component, and the bond coat includes a
chromide diffusion coating.
[0005] In some embodiments, the bond coat comprises a platinum
diffusion coating.
[0006] In some embodiments, the bond coat comprises an aluminide
diffusion coating.
[0007] In some embodiments, the bond coat further comprises
hafnium, silicon, zirconium, or any combination thereof.
[0008] In some embodiments, the bond coat is a diffusion coating
comprising chromium, platinum, aluminum, and at least one of
hafnium, silicon or zirconium.
[0009] In some embodiments, the base material is a Ni-based
superalloy having a gamma phase and a gamma prime phase distributed
within the gamma phase. Platinum of the diffusion coating resides
in the gamma prime phase, and chromium of the diffusion coating
resides in the gamma phase.
[0010] In some embodiments, the component is a gas turbine blade
comprising an airfoil that is exposed to combustion gases of a gas
turbine engine when in use. The bond coat and the thermal barrier
coating are located along the airfoil.
[0011] In accordance with various embodiments, a method of making a
gas turbine component having a thermal barrier coating defining at
least a portion of an outer surface of the gas turbine component
includes the step of forming a diffusion bond coat on a component
substrate. The diffusion bond coat includes chromium interdiffused
with a metal substrate material of the component substrate, and the
thermal barrier coating is subsequently coated over the bond
coat.
[0012] In some embodiments, the diffusion bond coat comprises a
Pt-aluminide coating.
[0013] In some embodiments, the step of forming the diffusion bond
coat comprises vapor phase deposition of the chromium on the metal
substrate material.
[0014] In some embodiments, the diffusion bond coat comprises
hafnium, silicon, zirconium, or any combination thereof.
[0015] In some embodiments, the step of forming the diffusion bond
coat comprises the steps of coating a slurry comprising a
platinum-group metal over the substrate material and heat treating
the slurry-coated substrate material to interdiffuse the
platinum-group metal with the substrate material.
[0016] In some embodiments, the slurry further comprises hafnium,
silicon, zirconium, or any combination thereof
[0017] In some embodiments, the step of forming the diffusion bond
coat further comprises vapor phase deposition of the chromium on
the metal substrate material before the step of coating the slurry
over the substrate material.
[0018] In some embodiments, the slurry comprises the chromium of
the bond coat, and the chromium is interdiffused with the substrate
material during the step of heat treating.
[0019] In some embodiments, the method includes vapor phase
aluminide coating the component substrate after the step of heat
treating, whereby the diffusion bond coat further comprises an
aluminide coating.
[0020] In some embodiments, the method includes forming a thermally
grown oxide layer over the diffusion bond coat, and the thermal
barrier coating is subsequently coated over the thermally grown
oxide layer.
[0021] In accordance with various embodiments, a gas turbine
component for use in a gas turbine engine includes a substrate
comprising a metal base material, a metal bond coat formed on the
base material, a thermally grown oxide layer formed on the bond
coat and comprising an oxide of a metal element of the bond coat,
and a ceramic-based thermal barrier coating disposed over the
substrate and defining at least a portion of an outer surface of
the gas turbine component. The thermal barrier coating further
comprises the oxide of the metal element of the bond coat.
[0022] In some embodiments, the oxide of the metal element is
alumina or chromia.
[0023] In some embodiments, the bond coat includes a diffusion
coating comprising chromide, Pt-aluminide, and at least one of
hafnium, silicon, or zirconium.
[0024] In accordance with various embodiments, a method of making a
gas turbine component includes the step of coating a turbine
component substrate with a coating system that includes a
ceramic-based thermal barrier coating having a sacrificial oxide
that is soluble in a molten sulfate salt.
[0025] It is contemplated that the various features set forth in
the preceding paragraphs, in the claims and/or in the following
description and drawings may be taken independently or in any
combination thereof. For example, features disclosed in connection
with one embodiment are applicable to all embodiments, except where
there is incompatibility of features.
BRIEF DESCRIPTION OF THE DRAWINGS
[0026] Illustrative embodiments will hereinafter be described in
conjunction with the appended drawings, wherein like designations
denote like elements, and wherein:
[0027] FIG. 1 is a perspective view of an exemplary gas turbine
component that may include the coating system described herein;
[0028] FIG. 2 is a cross-sectional view of a portion of a gas
turbine component illustrating an embodiment of the coating system
comprising a diffusion bond coat;
[0029] FIG. 3 is a cross-sectional view of a portion of a gas
turbine component illustrating an embodiment of the coating system
comprising an overlay bond coat;
[0030] FIG. 4 is a process flow diagram illustrating an exemplary
method of coating a gas turbine component substrate; and
[0031] FIG. 5 is a process flow diagram illustrating another
exemplary method of coating a gas turbine component substrate.
DETAILED DESCRIPTION
[0032] The coating system and methods described herein address
multiple problems that are likely to be encountered with turbine
components as operating temperatures increase, particularly in a
marine environment and particularly in applications in which the
turbine operates at or near peak power levels a majority of the
time.
[0033] In the marine environment, airborne sea salt can be ingested
by the turbine engine and, even though filtration systems may
reduce the level of sea salt, sufficient levels may be present to
condense sulfate salts onto some surfaces. The dew point for
condensation of sodium sulfate is dependent on pressure and salt
concentration but is generally below about 1500-1700.degree. F.
Conventional turbine materials with conventional hot corrosion
resistant coatings can survive for about 20,000 hours under such
conditions, even without a thermal barrier coating (TBC). For
next-generation gas turbines, one might expect that higher
operating temperatures would cause turbine components to remain
above the dew point of sodium sulfate such that less sodium sulfate
would condense and less sulfidation-type hot corrosion would occur,
leaving oxidation resistance as the main challenge.
[0034] However, it has been found that deposition of such salts
will occur during low power operation, when even TBC surfaces are
below the dew point of sulfate salts. When present on the surface
of a TBC, the molten salt can wick into the porosity of the coating
or into the columnar structure of certain TBCs (e.g., those applied
via electron beam physical vapor deposition) and penetrate to the
thermally grown oxide (TGO) layer on the underlying bond coat. Once
penetrated into the vacancies of the TBC, these salts may have a
very long residence time even when temperatures later rise above
their dew point. Thus, the rate of sulfidation-type corrosion
beneath the TBC can increase, and the TBC life to spallation may
drop to unacceptably low levels. When a portion of the TBC is lost
to spallation, component temperatures increase locally at the
spalling site, and hot corrosion/oxidation rates will be excessive,
leading to early engine removals.
[0035] Another sulfate salt that may be present in the marine
environment is calcium sulfate, the presence of which may modify
the corrosion environment and indirectly contribute to more severe
(e.g., acidic) conditions. This condition has been observed in
under-platform locations on turbine blades and may well become a
factor in future airfoil TBC applications. The coating system
described herein addresses this problem, among others.
[0036] Another failure mechanism in turbine components involves
impact damage at the leading edge of turbine blade airfoils. TBC's
are somewhat susceptible to local damage from large particle (50
.mu.m or larger) impact. This can lead to local loss of the TBC
accompanied by a local rise in underlying component temperature. In
aircraft engines, the time at maximum operating temperature is
relatively low because only a small percentage of overall operating
time (e.g., during aircraft acceleration for take-off) is at peak
power levels. In such cases, conventional bond coats can often
protect the component substrate at the local spall site for its
full service live. This will not be the case in aero-derivative gas
turbines that operate at full power most of the time. In these
applications, local spalls in the TBC can lead to burn through,
even with some level of de-rated power. Next-generation marine gas
turbines would thus benefit from improved TBC impact resistance
and/or a more graceful failure mode, as made possible by the
coating system described herein.
[0037] Further, there is a trend toward more reliance on the
coating system to protect the underlying substrate material, with
next-generation marine gas turbines likely using 2.sup.nd
generation single crystal alloys that have much lower hot corrosion
resistance than conventional superalloys (e.g., Rene 80). Should
the bond coat be penetrated, whether by impact damage or spallation
from corrosion beneath the TBC or some other cause, such substrates
may rapidly succumb to hot corrosion attack.
[0038] The coating system described below may be used on gas
turbine blades or other gas turbine components, such as compressor
blades, turbine or compressor vanes, seals, rotors or hubs, shafts,
shrouds, blisks, or any other component that may encounter the
harsh environment present in a gas turbine or other type of
combustion engine. The coating system includes a thermal barrier
coating (TBC) and an underlying bond coat that together function to
improve hot corrosion resistance of the coated component. The
coating system is particularly suitable for use in marine or other
applications where airborne salts are sometimes ingested by the
turbine where they react with other combustion constituents and/or
fuel contaminants and/or become molten. While described in the
context of a marine gas turbine engine intended to power waterborne
naval vessels, for example, the coating system may be employed in
other applications while realizing the same or other
advantages.
[0039] Referring to FIG. 1, an exemplary gas turbine component is
shown. In this embodiment, the component is a gas turbine blade 10.
The turbine blade 10 includes an airfoil 12 and a shank 14, each
extending from opposite sides of a platform 16. The airfoil 12 may
include a cross-section or profile configured to cause high and low
pressure regions on opposite sides thereof when placed in a
particular orientation in a flowing fluid, thus causing the blade
10 to move in the direction of lower pressure. The shank 14 may be
used as part of an attachment to secure the blade to a hub or other
component that rotates about a central axis. The shank 14 may
include several features not individually described here, such as a
root, neck, ridges, sealing flanges, "angel wings," etc. In
operation, multiple blades 10 may be arranged so that the airfoils
extend radially away from the central axis of the hub to form a
turbine that can transform energy from axial gas flow into
rotational motion, or vice versa where the blade is used as part of
a compressor or turbine.
[0040] The platform 16 lies between the airfoil 12 and the shank
14, generally dividing the blade into an upper boundary region 18
and a lower boundary region 20. The upper boundary region 18, also
referred to as the gas path region, is exposed to combustion gases
during operation and includes the airfoil 12 and a topside 22 of
the platform 16. The lower boundary region 20 is generally not
exposed to combustion gases during operation and includes an
underside 24 of the platform 16 and any other blade components
under the platform or on the shank side of the platform. This
arrangement of turbine blade components may also result in the
lower boundary region 20 operating at temperatures lower than those
at which the gas path region 18 operates. For example, the lower
boundary region 20 may operate at temperatures that range from
about 1200-1600.degree. F., while the gas path region 18 may
operate at higher temperatures that may range from about
1900-2100.degree. F. For purposes of this disclosure, the gas path
region 18 may also be referred to as the airfoil portion, and the
lower boundary region 20 may be referred to as the shank
portion.
[0041] In the illustrated embodiment, the turbine blade 10 also
includes internal cooling channels 26, the ends of some of which
are shown along the airfoil surface. Channels 26 may extend from
one or more surfaces of the shank portion 20 to one or more
surfaces of the airfoil portion 18 to facilitate the flow of a
cooling fluid such as air therethrough. Various blade cooling
arrangements are known in the art, and the cooling channels may be
omitted entirely in some cases.
[0042] Due to the harsh environment in and around an operating gas
turbine engine, engine components are sometimes constructed using
superalloy materials that have high strength, ductility, and creep
resistance at high temperatures and relatively high resistance to
corrosion. Superalloy materials may be based on nickel (Ni), cobalt
(Co), and/or Iron (Fe). Examples of superalloys include alloys
available under the trade names Hastelloy, Inconel, and Rene, such
as Rene N4, Rene N5, or others. While the corrosion-resistance of
superalloys may generally be considered very good as metal alloys
are concerned, the elevated temperatures and stresses, corrosive
combustion gases, and other elements (e.g., atmospheric pollutants
or particulates, fuel additives and impurities, salts, etc.) in the
gas turbine operating environment can accelerate the corrosion of
even the most corrosion-resistant superalloys. As used herein,
corrosion may be divided into two categories, including Type I and
Type II corrosion. Type I corrosion generally refers to relatively
high temperature corrosion that occurs above about 1600.degree. F.
and includes sulfidation-type corrosion. Type II corrosion
generally refers to relatively low temperature corrosion that
occurs below about 1600.degree. F.
[0043] Various types of coatings or surface treatments have been
developed in attempt to improve the corrosion and fatigue
resistance of superalloy components in gas turbines. One approach
includes the use of a thermal barrier coating (TBC), particularly
along the airfoil portion of turbine blades. TBCs are generally
designed to thermally insulate the underlying materials from the
high temperature combustion gases so that less thermal energy is
absorbed by the turbine blade, thereby allowing the superalloy or
other substrate material to operate at a lower temperature than
would otherwise be possible. In some cases, the gas turbine
component can operate in an environment in which the temperature of
the combustion gases exceeds the melting point of the substrate
material when a suitable TBC is employed.
[0044] Referring to FIGS. 2 and 3, partial cross-sections of
exemplary gas turbine components 10 are shown, including a gas
turbine component substrate 30, such as a blade substrate, and a
coating system 40 that includes a thermal barrier coating 42. The
substrate 30 may be formed from a base material 32 by casting
and/or other known processing techniques. The base material 32 may
be a metal or metal-based alloy capable of forming a gamma/gamma
prime (.gamma./.gamma.') microstructure in which the gamma prime
phase is in the form of a precipitate distributed within the gamma
phase matrix. One example of such an alloy is a Ni-based superalloy
that can form a gamma/gamma prime microstructure when heat-treated
under certain conditions. For example, heat-treating a Ni-based
superalloy can cause a gamma prime phase to form as a precipitate
that includes Ni.sub.3Al and/or Ni.sub.3Ti distributed in a gamma
phase that is a solid solution including Ni and other elements. Any
of the above-mentioned exemplary superalloys, as well as other
alloys capable of forming a gamma/gamma prime microstructure, may
be suitable for use as the base material 32. Cobalt-based
superalloys may also be suitable. The substrate 30 may also include
materials or layers of materials other than the base material 32.
For example, substrate 30 may include a layer of the base material
32 clad or otherwise attached to a different underlying
material.
[0045] The coating system 40 includes the thermal barrier coating
42 along with one or more other coatings or material layers that
overlie the base material 32 and provides an increased resistance
to corrosion relative to the base material, particularly at high
temperatures and/or in an operating environment that includes
airborne salts, such as sulfate salts. The coating system 40 of
FIG. 2 includes a diffusion bond coat 44 at an outer surface 46 of
the substrate 30 and an oxide layer 48 between the bond coat 44 and
the TBC 42. The coating system 40 of FIG. 3 includes the same
layers as the system of FIG. 2, except the bond coat 44' is an
overlay coating rather than a diffusion coating.
[0046] Each coating layer may be classified as either an overlay
coating or a diffusion coating. Both types of coatings may be at
least partially interdiffused with an underlying material, but any
interdiffusion that is present with an overlay coating is in the
form of a relatively thin layer at the interface of the overlay
coating and the underlying material. A diffusion coating has a
substantial portion of its thickness interdiffused with the
underlying material, and may be entirely interdiffused with the
underlying material. Diffusion coatings may be designated with an
"-ide" suffix and indicate the element or elements interdiffused
with the underlying material. For example, aluminide and chromide
coatings are diffusion coatings of aluminum and chromium,
respectively. Thus, a diffusion coating is a layer of material that
is richer in the coated and diffused element(s) than the underlying
material and further includes the constituent elements of the
underlying coated material.
[0047] The thermal barrier coating 42 is an overlay coating and, as
noted above, is generally configured to thermally insulate the
substrate 30 from the high combustion gas temperatures in
operation. Suitable thermal barrier coatings may be ceramic-based
materials, such as yttria-stabilized zirconia (YSZ) or a rare earth
zirconate. The TBC may be coated over the substrate 30 and bond
coat by known methods.
[0048] The TBC 42 may be porous, have a specific microstructure
(e.g., columnar voids), or have cracks or microcracks oriented
perpendicular to the outer surface of the TBC and thereby allow
corrosive elements such as oxygen or other substances to migrate
through the TBC toward the substrate 30. One of the functions of
the bond coat 44 is to provide corrosion protection for the
substrate material 32. Corrosion of the substrate material 32
beneath the TBC 42 can be doubly problematic because it not only
affects the integrity of the substrate material, but it can also
cause spalling of the TBC due to the stresses induced in the TBC by
the growth of the underlying corrosion. When TBC spalling occurs,
the thermal barrier properties of the TBC are lost locally and the
underlying material is exposed to the high temperature combustion
gases, accelerating any corrosion and reducing fatigue life.
[0049] The oxide layer 48 is not necessarily present during turbine
blade fabrication. The oxide layer 48 may be a thermally grown
oxide (TGO) that forms above about 1300.degree. F. from oxidation
of the material (e.g., the bond coat 44) beneath the TBC 42,
through which oxygen and other gases passes through during
operation due to TBC porosity. In some embodiments of the coating
system 40 described herein, the oxide layer 48 is pre-grown or
otherwise formed prior to application of the TBC 42. As discussed
below, controlling the composition of the oxide layer 48 has been
found to be advantageous, particularly with respect to
sulfidation-type corrosion. In addition, the TGO has been found to
greatly enhance adhesion of the TBC to the bond coat 44.
[0050] In the embodiment of FIG. 2, the bond coat 44 may include a
chromide diffusion coating. While chromide diffusion coatings have
been proposed for use in turbine components to improve corrosion
resistance, applications have been limited to lower temperature
portions of such components, such as the shank portion of a turbine
blade. Chromide diffusion coatings have generally been considered
insufficient on higher temperature regions of turbine components,
such as the airfoil portion of a turbine blade. However, inclusion
of chromide in the bond coat 44 has now been found to offer certain
advantages. In particular, the presence of chromide in the bond
coat 44 ensures the presence of chromia in the oxide layer 48,
which behaves differently from the alumina that is present in the
oxide layer when aluminide is present in the bond coat,
particularly with respect to sulfidation-type corrosion.
[0051] For example, it has been determined that a molten film of
sodium sulfate on a substrate surface without a TBC will flux
(i.e., dissolve) a protective alumina film formed thereon. However,
alumina does not saturate the molten salt in a manner that prevents
further fluxing. Instead, the alumina (or aluminate) reprecipitates
near the outer surface of the film of molten salt, and fluxing
continues. Chromia behaves differently. While chromia may also
dissolve in the molten salt, it reprecipitates at the lower
oxygen-activity location within the molten salt--i.e., near the
protective oxide layer. In this manner, the presence of chromia in
the oxide layer 48 of the illustrative coating system 40 offers
better Type I corrosion protection and, in particular, better
resistance to sulfidation-type corrosion. The presence of the
chromide diffusion coating in the bond coat 44 can help ensure the
presence of chromia in the TGO 48 and has thus unexpectedly been
found to offer improved high temperature corrosion resistance when
used as part of the bond coat 44 beneath the TBC 42.
[0052] The chromide diffusion coating can be formed in a number of
ways, each including deposition of chromium on the substrate
surface to be coated and interdiffusion of the chromium with the
substrate material by heat treatment. In one embodiment, the
chromide diffusion coating is formed by vapor phase deposition, in
which the substrate is exposed to a vapor containing the chromium
to be deposited. Vapor phase deposition is typically performed in a
chamber environment at a temperature sufficient to diffuse the
chrome into the substrate surface immediately upon deposition such
that separate deposition and heat treatment steps are not required.
For example, the same elevated temperature required for diffusion
can be used to heat a bed containing chromium metal and an
activator such that chromium-containing vapor is produced, with the
vapor coming into contact with the substrate and depositing the
chromium on the substrate, which may be supported over the bed
vapor source. The chromide diffusion coating may also be formed in
a slurry process, in which the chromium is deposited on the
substrate surface as part of a coating comprising the chromium, an
activator, and a liquid carrier that allows the coating to be
coated onto the substrate by spraying, painting, dipping, etc.,
with the liquid component being evaporated away prior to diffusion
by heat treating. The chromide diffusion coating may also be formed
in a pack cementation process, in which the chromium is deposited
on the substrate surface by packing the substrate surface in a bed
(e.g., a powder bed) of material containing chromium and activator,
similar to the above-described vapor phase deposition bed. In both
the pack cementation and slurry processes, chromium-containing
vapor may be produced from the respective bed of material and
slurry coating during the heat treatment step. But these processes
differ from vapor phase deposition in that at least some of the
chromium metal is in contact with the substrate before heat
treating begins and during heat treating. Also, other constituents
can be added to the slurry coating mixture or to the
pack-cementation bed to be diffused into the substrate. Other
deposition methods can be used as well, such as electroplating,
PVD, or other suitable techniques, followed by diffusion by heat
treatment.
[0053] In one embodiment, the chromide diffusion coating is formed
by combining a powder contact process and a simultaneous vapor
phase process, resulting in a super-chromide diffusion coating, at
least a portion of which has a chromium content of 60 wt % or
higher. In one example, the portion of the substrate to be coated
is packed into a bed of powder material comprising chromium metal
and an activator. Then, while heating the substrate in the bed of
material, a chromium-containing vapor is flowed through the bed of
material from a vapor source outside of the bed of material to
boost the concentration of chrome-containing vapor that the packed
portion of the substrate is exposed to. The result is a chromide
diffusion coating with a higher concentration of chromium than can
be produced by any other chromide coating technique alone.
Super-chromide coatings and methods of producing them are discussed
further in U.S. Patent Application Publication No. 2013/0115907,
hereby incorporated by reference.
[0054] Where the bond coat 44 is a diffusion coating as illustrated
in FIG. 2, it may also include one or more of the following
elements interdiffused with the substrate material: a
platinum-group metal such as platinum (Pt), aluminum (Al), hafnium
(Hf), silicon (Si), or zirconium (Zr).
[0055] Platinum and aluminum can be combined in diffusion coatings
in the form of a bond coat as a Pt-aluminide coating.
Conventionally, a Pt-aluminide bond coat is formed on a substrate
by first electroplating the Pt onto the substrate surface, then
using vapor phase deposition to deposit the aluminum on the part
surface while the high temperature environment of the vapor phase
process diffuses both the Pt and the aluminum into the substrate
surface to form the Pt-aluminide diffusion coating. Alternatively,
the electroplated Pt may be heat treated in a vacuum to diffuse it
into the substrate prior to the vapor phase aluminide process.
[0056] Where the substrate material is a metal or metal-based alloy
capable of forming a gamma/gamma prime (.gamma./.gamma.')
microstructure as described above, the Pt or other platinum-group
metal can be interdiffused with the substrate material in a
separate step from the aluminide coating to form a gamma/gamma
prime diffusion coating. In this type of diffusion coating, the
Pt-group metal resides in the gamma prime phase of the heat treated
component. Such coatings have recently found uses as
corrosion-resistant coatings separately from their utility as bond
coats for TBCs. For example, U.S. Pat. No. 9,297,089, incorporated
herein by reference, describes such coatings used on the lower
temperature shank portion of turbine blades in the absence of
aluminide and TBCs. This type of coating has also been found to
perform well as a bond coat for a TBC, but it can be susceptible to
hot corrosion and oxidation failure quickly if the TBC coating is
compromised.
[0057] In one embodiment, the bond coat 44 comprises a
platinum-group diffusion coating. In a particular embodiment, the
platinum-group diffusion coating is formed by a slurry process
similar to that described above in conjunction with the diffusion
chromide coating. For example, a platinum-group metal such as Pt
may be deposited on the substrate surface as part of a slurry
coating comprising the Pt-group metal, an activator, and a liquid
carrier that allows the coating to be coated onto the substrate by
spraying, painting, dipping, etc., with the liquid component being
evaporated away prior to diffusion by heat treating. Application by
the slurry process offers certain advantages over electroplating,
which has long been the preferred method of applying Pt in a
Pt-aluminide coating. For example, other elements can be added to
the slurry mixture to be simultaneously applied and subsequently
diffused into the substrate material together with the Pt-group
metal. In a particular embodiment, the above-described chromide
diffusion coating is formed in the same slurry process as the
Pt-group diffusion coating by including both chromium and the
Pt-group metal in the slurry mixture. The slurry process also
facilitates selective deposition of the coating constituents. For
example, the slurry can be sprayed or otherwise coated onto the
turbine component substrate only where desired, while
electroplating typically results in the deposition of platinum over
the entire conductive substrate surface, thereby wasting an
expensive coating constituent.
[0058] In other embodiments, a dopant is added to the Pt-group
slurry, such as hafnium (Hf), silicon (Si), or zirconium (Zr).
These dopants are diffused into the substrate material as part of
the bond coat 44 and act to stabilize the TGO layer 48 to enhance
TBC adherence.
[0059] The bond coat 44 may further include an aluminide diffusion
coating, which may be formed via vapor phase deposition and
simultaneous heat treatment to diffusion the aluminum into the
substrate material. When the aluminide diffusion coating is formed
after the Pt-group metal is already interdiffused with the
substrate material, the result is a complex Pt-aluminide coating
that is somewhat different from a Pt-aluminide coating that is
formed when electroplated Pt and aluminum are simultaneously
diffused into the substrate material. For example, the pre-diffused
Pt-group metal already resides in a gamma prime phase of the
coating before the aluminum is diffused into the already existing
coating. The aluminide forms a particularly complex Pt-aluminide
when the Pt-group diffusion coating is formed via a slurry process
in which the slurry includes a TGO stabilizing dopant (e.g.,
hafnium).
[0060] FIG. 4 illustrates an exemplary method 100 of forming the
coating system according to a specific embodiment. The illustrated
method includes the step 110 of providing a turbine component
substrate, the step 120 of forming a diffusion bond coat over at
least a portion of the substrate, and the step 130 of forming a
thermal barrier coating that provides the outer surface of the
finished turbine component. In this example, the diffusion bond
coat is formed via steps 140, 150, and 160, including respectively
forming a chromide diffusion coating, a doped Pt-group slurry
diffusion coating, and an aluminide diffusion coating on the
substrate. Step 150 is performed via a slurry process as described
above in which the slurry mixture includes the Pt-group metal
(e.g., platinum) along with one or more TGO-stabilizing dopants,
such as Hf, Si, or Zr. Each of steps 140 and 160 may be performed
via a vapor phase deposition process as described above or by other
suitable processes such as a pack cementation process or separate
slurry processes. In one specific embodiment, step 150 includes
step 140 with the chromium being included as part of the doped-Pt
slurry mixture.
[0061] In another specific embodiment, step 140 is performed in a
combined process in which the chromium is provided by two different
sources, including a powder source in contact with the substrate
(e.g., pack cementation) and a vapor source separate from the
powder source (e.g., chrome-containing vapor flowed through the
pack cementation bed), resulting in a super-chrome diffusion
coating before the slurry coat and aluminide coating. The doped
Pt-group diffusion coating may then be formed by applying a
dopant-containing Pt-slurry over the super-chrome coated substrate
and diffusing the Pt-group metal and dopants into the chrome-coated
surface. This results in a gamma/gamma prime system in which the
Pt-group metal resides in the gamma prime phase of the heat treated
substrate material with the chrome residing in the gamma phase
(e.g., with the nickel component of the substrate material in solid
solution). A vapor phase aluminide diffusion coating is then formed
over the gamma/gamma prime system, resulting in a complex
Pt-aluminide coating. After completion of the step 120 of forming
the bond coat, the TBC is applied thereover in step 130. The TBC
may be a ceramic or ceramic-based coating such as yttria-stabilized
zirconia (YSZ) or other suitable material and may be applied via an
air plasma process, solution plasma spray (SPS), electron-beam
physical vapor deposition (EB-PVD), or other suitable process.
[0062] The complex Pt-aluminide coating thus formed results in the
dopants being located in specific locations of the bond coat where
it can be beneficial to the TGO and/or TBC layers. More
particularly, the dopants (e.g., Hf, Si, and/or Zr) will be
distributed throughout the thickness of the aluminide portion of
the bond coat, with some of the dopant near the aluminide-coated
surface and some of the dopant diffused farther into the substrate.
In contrast, adding such dopants during the vapor phase aluminide
process after more conventional Pt electroplating may result in
aggregation of the dopants at the outermost portion of the
aluminide as a localized hafnide that is detrimental to formation
of the TGO layer where present at the bond coat surface, which
negatively affects TBC adhesion. In order to be beneficial, dopants
such as hafnium must be present in lower concentrations and more
evenly distributed within the thickness of the bond coat. It is
desirable to have the dopants a few microns (e.g., 1-3 .mu.m) from
the TGO and below the surface of the bond coat. This is very
difficult to achieve with Pt electroplate followed by a doped
aluminide vapor phase process.
[0063] The example of FIG. 4 also illustrates an optional step 170
in which a thermally grown oxide layer is formed on the bond coat
prior to application of the thermal barrier coating. In another
example, the TGO layer can be formed after application of the TBC
but before the finished turbine component is put into service. In
either case, pre-forming the TGO layer offers the advantage of
forming the oxide under controlled conditions rather than allowing
it to form under the unpredictable conditions present during
initial operation of the component.
[0064] FIG. 5 illustrates another exemplary method 200 of forming
the coating system according to a specific embodiment. The
illustrated method includes the step 210 of providing a turbine
component substrate, the step 220 of forming a bond coat over at
least a portion of the substrate, and the step 230 of forming a
doped thermal barrier coating that provides the outer surface of
the finished turbine component. In this example, the diffusion bond
coat is formed via step 250 of forming an overlay bond coat (e.g.,
MXCrAly) by known methods. Optionally, a chrome diffusion coating
is formed in step 240 prior to application of the overlay bond
coat.
[0065] In this example, the TBC is doped to include a sacrificial
oxide which will dissolve in and at least partially saturate any
molten salts, such as sodium sulfate, that makes its way into the
porosity of the TBC. Suitable dopants include alumina or chromia
which will dissolve in the molten salt and thus slow or halt
fluxing of the TGO layer. The dopant may be applied with the TBC in
a thermal spray process in an amount of about 1 wt % to 3 wt %. The
example of FIG. 5 also illustrates an optional step 270 in which
the TGO layer is formed on the bond coat prior to application of
the thermal barrier coating. In another example, the TGO layer can
be formed after application of the TBC but before the finished
turbine component is put into service.
[0066] It is to be understood that the foregoing is a description
of one or more preferred exemplary embodiments of the invention.
The invention is not limited to the particular embodiment(s)
disclosed herein, but rather is defined solely by the claims below.
Furthermore, the statements contained in the foregoing description
relate to particular embodiments and are not to be construed as
limitations on the scope of the invention or on the definition of
terms used in the claims, except where a term or phrase is
expressly defined above. Various other embodiments and various
changes and modifications to the disclosed embodiment(s) will
become apparent to those skilled in the art. All such other
embodiments, changes, and modifications are intended to come within
the scope of the appended claims.
[0067] As used in this specification and claims, the terms "for
example," "e.g.," "for instance," "such as," and "like," and the
verbs "comprising," "having," "including," and their other verb
forms, when used in conjunction with a listing of one or more
components or other items, are each to be construed as open-ended,
meaning that that the listing is not to be considered as excluding
other, additional components or items. Other terms are to be
construed using their broadest reasonable meaning unless they are
used in a context that requires a different interpretation.
* * * * *