U.S. patent application number 10/712821 was filed with the patent office on 2005-05-12 for thermal barrier coating having a heat radiation absorbing topcoat.
This patent application is currently assigned to General Electric Company. Invention is credited to Ackerman, John Frederick, Ivkovich, Daniel Peter JR., Murphy, Jane Ann, Nagaraj, Bangalore Aswatha, Skoog, Andrew Jay, Stowell, William Randolph.
Application Number | 20050100757 10/712821 |
Document ID | / |
Family ID | 34435671 |
Filed Date | 2005-05-12 |
United States Patent
Application |
20050100757 |
Kind Code |
A1 |
Stowell, William Randolph ;
et al. |
May 12, 2005 |
Thermal barrier coating having a heat radiation absorbing
topcoat
Abstract
Coating systems suitable for protecting components exposed to
high-temperature environments, such as the hot gas flow path
through a gas turbine engine. A multilayer thermal barrier coating
(TBC) system characterized by a low coefficient of thermal
conductivity and having a heat-absorbing topcoat comprised of the
thermal decomposition product of at least one metallic element and
at least one ceramic precursor binder component.
Inventors: |
Stowell, William Randolph;
(Rising Sun, IN) ; Nagaraj, Bangalore Aswatha;
(West Chester, OH) ; Skoog, Andrew Jay; (West
Chester, OH) ; Murphy, Jane Ann; (Franklin, OH)
; Ackerman, John Frederick; (Laramie, WY) ;
Ivkovich, Daniel Peter JR.; (Fairfield, OH) |
Correspondence
Address: |
MCNEES WALLACE & NURICK LLC
100 PINE STREET
P.O. BOX 1166
HARRISBURG
PA
17108-1166
US
|
Assignee: |
General Electric Company
Schenectady
NY
|
Family ID: |
34435671 |
Appl. No.: |
10/712821 |
Filed: |
November 12, 2003 |
Current U.S.
Class: |
428/621 ;
428/624; 428/632 |
Current CPC
Class: |
Y02T 50/67 20130101;
Y10T 428/12611 20150115; F05D 2300/2118 20130101; Y10T 428/12535
20150115; C23C 24/08 20130101; Y10T 428/12556 20150115; C23C
28/3455 20130101; F01D 5/288 20130101; Y02T 50/60 20130101; C23C
28/3215 20130101; Y02T 50/671 20130101; Y02T 50/6765 20180501; C23C
28/321 20130101; C23C 28/324 20130101; C23C 8/42 20130101; C23C
8/80 20130101; C23C 28/345 20130101 |
Class at
Publication: |
428/621 ;
428/624; 428/632 |
International
Class: |
B21D 039/00 |
Claims
1. A superalloy component comprising: a superalloy substrate; an
alumina-containing bond coat layer on the substrate; an adherent
layer of ceramic material forming a thermal barrier coating on the
alumina-containing bond coat layer; and a heat-absorbing topcoat
layer applied to the thermal barrier coating, the heat-absorbing
topcoat layer comprised of the thermal decomposition product of a
mixture comprised of at least one metallic element and at least one
ceramic precursor binder component.
2. An article according to claim 1, wherein the topcoat layer has a
thickness of between about 2 to about 30 mils.
3. An article according to claim 2, wherein the topcoat layer has a
thickness of between about 8 to about 12 mils.
4. An article according to claim 1, wherein the at least one
metallic element is selected from the group consisting of alumina,
hafnia, tantala, silica, platinum, nickel, iron, cobalt, chromium
oxide, and rare earth oxides
5. An article according to claim 1, wherein the at least one
ceramic precursor binder component is selected from the group
consisting of silicone, alumoxanes, plasticized titanium ethoxide
or tantalum ethoxide, and other plasticizable metal organic
compounds.
6. An article according to claim 1, wherein upon application and
prior to thermal decomposition the topcoat layer is comprised of at
least 25 weight percent metallic element, and at least 25 weight
percent of the ceramic precursor binder.
7. An article according to claim 1, wherein the adherent layer of
ceramic material is comprised of between about 3% to about 20%
yttrium-stabilized zirconia (YSZ).
8. An article according to claim 1, wherein the overcoat layer is
applied by spraying a mixture of the metallic element and the
ceramic precursor binder onto the thermal barrier coating followed
by heating of the deposited mixture to thermally convert the binder
and to oxidize at least a portion of the metallic element to form a
continuous topcoat layer.
9. An article according to claim 1, wherein the overcoat layer
further comprises at least two thin layers of heat-absorbing
material.
10. An article according to claim 9, wherein each thin layer is
comprised of a different metallic element.
11. A superalloy article having a thermal barrier coating system,
the system comprised of an alumina-containing bond coat layer, an
adherent layer of ceramic material forming a thermal barrier
coating on the alumina-containing bond coat layer; and a
heat-absorbing topcoat layer applied to the thermal barrier
coating, the heat-absorbing topcoat layer comprised of the thermal
decomposition product of a mixture comprised of at least one
metallic element and at least one ceramic precursor organic binder
component, the article made in accordance with the method
comprising the steps of: preparing an exposed surface of the
thermal barrier coating in order to provide a predetermined
roughness of the exposed surface and to remove debris and
contaminants from the exposed surface; and applying a
heat-absorbing topcoat layer onto the exposed surface so as to
overlie the thermal barrier coating and form a continuous overcoat
on the thermal barrier coating.
12. The article according to claim 11, wherein the predetermined
roughness is between about 80-125 microinches.
13. The article according to claim 11, wherein the step of
preparing the exposed surface comprises measuring the roughness of
the exposed surface; determining the difference between the
predetermined roughness and the measured roughness; and polishing
or roughening the exposed surface based upon the difference until
the roughness is about equal to the predetermined roughness.
14. The article according to claim 11, wherein the heat-absorbing
overcoat is comprised of a mixture comprising a ceramic precursor
in at least 25 weight percent, and a metallic element in at least
25 weight percent.
15. The article according to claim 11, wherein the step of applying
is performed using a process selected from the group consisting of
spraying, chemical vapor deposition, physical vapor deposition,
plasma spraying, and sputtering.
16. The article according to claim 11, wherein step of applying is
performed until the overcoat is between about 2 to about 30 mils
thick.
17. The article according to claim 11, wherein step of applying is
performed until the overcoat is between about 8 to about 12 mils
thick.
18. An article according to claim 11, wherein the overcoat layer
further comprises at least two thin layers of heat-absorbing
material.
19. An article according to claim 18, wherein each separate layer
is comprised of a different metallic element.
20. A coated superalloy component for use in the flowpath of a gas
turbine engine, the coated component made by the process of:
providing a superalloy component comprising a flowpath part from a
gas turbine engine assembly; applying a thermal barrier coating
system, the system comprised of an alumina-containing bond coat
layer, an adherent layer of ceramic material forming a thermal
barrier coating on the alumina-containing bond coat layer; and
applying a heat-absorbing topcoat layer to the exposed surface of
the thermal barrier coating.
21. The coated article made by the process of claim 20, wherein the
step of applying a heat-absorbing topcoat layer is comprised of the
steps of: providing a mixture comprised of at least one metallic
element and at least one ceramic precursor binder component;
spraying the mixture onto the exposed surface of the thermal
barrier coating so as to overlie the thermal barrier coating and
form a continuous topcoat on the thermal barrier coating; and
heating the topcoat to a degree sufficient to thermally convert the
at least one ceramic precursor binder component and to oxidize at
least a portion of the metallic element to yield a heat-absorbing
ceramic-metallic matrix topcoat.
22. The coated article made by the process of claim 20, wherein the
step of applying a heat-absorbing topcoat layer is comprised of the
steps of: providing at least one metallic element selected from the
group consisting of alumina, hafnia, tantala, silica, platinum,
nickel, iron, cobalt, chromium oxide, and rare earth oxides; and
depositing the at least one metallic element onto the exposed
surface of the thermal barrier coating so as to overlie the thermal
barrier coating to form a heat-absorbing topcoat.
Description
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH
[0001] Not applicable.
FIELD OF THE INVENTION
[0002] This invention relates to coating systems suitable for
protecting components exposed to high-temperature environments,
such as the hot gas flow path through a gas turbine engine. More
particularly, this invention is directed to a multilayer thermal
barrier coating (TBC) system characterized by a low coefficient of
thermal conductivity and having a heat-absorbing topcoat.
BACKGROUND OF THE INVENTION
[0003] The use of thermal barrier coatings (TBC) on components such
as combustors, high pressure turbine (HPT) blades and vanes is
increasing in commercial as well as military gas turbine engines.
The thermal insulation of a TBC enables such components to survive
higher operating temperatures, increases component durability, and
improves engine reliability. TBC is typically a ceramic material
deposited on an environmentally-protective bond coat to form what
is termed a TBC system. Bond coat materials widely used in TBC
systems include oxidation-resistant overlay coatings such as MCrAlX
(where M is iron, cobalt and/or nickel, and X is yttrium or another
rare earth element), pure oxidation resistant intermetallic
compounds such as nickel aluminides, and oxidation-resistant
diffusion coatings such as diffusion aluminides that contain
aluminum intermetallics.
[0004] Ceramic materials and particularly binary yttria-stabilized
zirconia (YSZ) are widely used as TBC materials because of their
high temperature capability, low thermal conductivity, and relative
ease of deposition such as by air plasma spraying (APS), flame
spraying such as hyper-velocity oxy-fuel (HVOF), physical vapor
deposition (PVD) and other known TBC application techniques. TBCs
formed by these methods generally have a lower thermal conductivity
than a dense ceramic of the same composition as a result of the
presence of microstructural defects and pores at and between grain
boundaries of the TBC microstructure.
[0005] TBC employed in the highest temperature regions of gas
turbine engines are often deposited by electron beam physical vapor
deposition (EBPVD), which yields a columnar, strain-tolerant grain
structure that is able to expand and contract without causing
damaging stresses that lead to spallation. Similar columnar
microstructures can be produced using other atomic and molecular
vapor processes, such as sputtering (e.g., high and low pressure,
standard or collimated plume), ion plasma deposition, and all forms
of melting and evaporation deposition processes (e.g., cathodic
arc, laser melting, etc.).
[0006] Regardless of the method of application, in order for a TBC
to remain effective throughout the planned life cycle of the
component it protects, it is important that the TBC has and
maintains a low thermal conductivity throughout the life of the
component, including high temperature excursions. However, the
thermal conductivities of TBC materials such as YSZ are known to
increase over time when subjected to the operating environment of a
gas turbine engine. As a result, TBCs for gas turbine engine
components are often deposited to a greater thickness than would
otherwise be necessary. Alternatively, internally cooled components
such as blades and nozzles must be designed to have higher cooling
flow. Both of these solutions are undesirable for reasons relating
to cost, component life and engine efficiency.
[0007] In view of the above, it can be appreciated that further
improvements in TBC technology are desirable, particularly as TBCs
are employed to thermally insulate components intended for more
demanding engine designs. For example, known TBCs applied by APS
methods at a thickness of about 12 mils exhibit heat radiation
transparency of about 40% in the wavelength range of about 0.5 to
about 5 microns. When exposed to a 4500R flame (R being the
absolute scale for Fahrenheit temperature scale, wherein absolute
zero is equal to 0 degrees R, and wherein absolute zero is equal to
-459.6 degrees F.), approximately 40% of the heat energy from the
flame in the 0.5 to 5 micron wavelength region is transmitted
through the TBC to the substrate despite the low thermal
conductivity of the TBC. This is illustrated by the FIG. 1, which
represents a chart showing measured transmittance of a typical 12
mil thick TBC coating.
[0008] From the below graph, one can see that a black body
radiating at .about.2060.degree. F. produces much less energy than
one radiating at 3680.degree. F. or 4220.degree. F. Thus, if a
surface is maintained at about 2060.degree. F. by means of
convection, it can only radiate a maximum energy of the lowest line
on the attached graph. However, in gas turbine combustors, for
example, gas temperatures are in the range of 3500.degree. F. to
4000.degree. F. and higher under maximum conditions, such as those
used for take-off of a gas turbine powered aircraft. If a TBC on
such a combustor is 40% transparent to heat in the 0.8 to 3 micron
range, much more energy will be transmitted through it from a
3680.degree. F. source than from a 2060.degree. F. source. Hence,
stopping heat radiation from a 3680.degree. F. source in an
absorbing media on the surface of the TBC and maintaining the
surface temperature of the TBC by means of convective airflow would
significantly reduce the energy that reaches the combustor's metal
wall when compared to allowing the heat from combustion flames to
directly impinge upon a semi-transparent TBC.
[0009] Reducing the total thermal load of the substrate would
reduce the part temperature, ceteris paribus. Such a reduction in
thermal load can be accomplished by providing a TBC coating system
with low thermal conductivity and having a heat-absorbing topcoat
that reduces the heat transparency of the coating system, thereby
reducing substrate component surface temperatures. Such a system
also allows for reduction in TBC coating thickness while
maintaining component surface temperature below a predetermined
maximum component surface temperature. Reduced TBC thickness,
especially in applications such as combustors that require
relatively thick TBCs, would result in a significant cost
reduction, as well as weight reduction benefits.
BRIEF SUMMARY OF THE INVENTION
[0010] The present invention provides a novel thermal barrier
coating (TBC) system having a heat-absorbing topcoat disposed over
a TBC undercoat, the topcoat reducing the thermal transparency of
the TBC system, resulting in low thermal conductivity and low
thermal transparency of the TBC system to the underlying substrate.
The TBC system may include a bond coat over the substrate, the bond
coat providing the means by which a TBC undercoat is adhered to a
component surface.
[0011] The TBC undercoat of this invention preferably comprises
yttria-stabilized zirconia (YSZ). An exemplary YSZ is described in
U.S. Pat. No. 6,586,115 "Yttria-stabilized zirconia with reduced
thermal conductivity." However, other zirconia-based ceramic
materials could also be used as the undercoat in accordance with
this invention, such as zirconia fully stabilized by yttria,
nonstabilized zirconia, or zirconia partially or fully stabilized
by ceria, magnesia, scandia and/or other oxides.
[0012] The topcoat may comprise any material that exhibits heat
adsorption in a desired spectrum at environmental operating
temperatures, such as but not limited to environmental operating
temperature of about 2500 to about 4500 degrees Fahrenheit.
Exemplary topcoat materials may comprise alumina, hafnia, tantala,
silica, platinum, nickel, iron, cobalt, chromium oxide, rare earth
oxides, silicides, carbides, and any combination thereof.
[0013] Another advantage of the topcoat is that it can be applied
by any of several methods. For example, it can be applied by
spraying a liquid mixture of at least one metallic element and at
least one a ceramic precursor binder, followed by heating to
thermally convert the binder to form a heat-absorbing
ceramic-metallic matrix coating containing the at least one
metallic element. Alternatively, the topcoat can be applied over a
TBC in the form of one or more thin film layers by methods such as,
but not limited to, chemical vapor deposition ("CVD"), physical
vapor deposition (PVD), plasma spray, and sputtering.
[0014] Other objects and advantages of this invention will be
better appreciated from the following detailed description.
BRIEF DESCRIPTION OF THE DRAWINGS
[0015] FIG. 1 represents a chart showing measured transmittance of
a typical 12 mil thick TBC coating.
[0016] FIGS. 2 through 3 represent cross-sectional views of thermal
barrier coating systems in accordance with the embodiments of the
present invention.
DETAILED DESCRIPTION OF THE INVENTION
[0017] The present invention is generally applicable to components
subjected to high temperatures, and particularly to components such
as the high pressure and low pressure turbine vanes (nozzles) and
blades (buckets), shrouds, combustor liners and augmentor hardware
of gas turbine engines. The invention provides a thermal barrier
coating (TBC) system suitable for protecting those surfaces of a
gas turbine engine component that are subjected to hot combustion
gases. While the advantages of this invention will be described
with reference to gas turbine engine components, the teachings of
the invention are generally applicable to any component on which a
TBC may be used to protect the component from a high temperature
environment.
[0018] The coating systems in accordance with the preferred
embodiments of this invention are represented in FIGS. 2 through 3.
In each embodiment, the coating system 10, 110 is shown as
including a metallic bond coat 12 that overlies the surface of a
substrate 14, the latter of which is typically a superalloy and the
base material of the component protected by the coating systems 10,
110. As is typical with TBC coating systems for gas turbine engine
components, the bond coat 12 is preferably an aluminum-rich
composition, such as an overlay coating of an MCrAlX alloy or a
diffusion coating such as a diffusion aluminide or a diffusion
platinum aluminide of a type known in the art. Aluminum-rich bond
coats of this type develop an aluminum oxide (alumina) scale, which
is grown by oxidation of the bond coat 12. The alumina scale 16
chemically bonds a TBC undercoating 18 to the bond coat 12 and
substrate 14. Preferably although not necessarily, the bond coat
layer 12 thickness is between about 2 to 10 mils and is uniform.
The bond coat layer 12 may also be treated, e.g., by peening and/or
heat treating, to densify any voids and to improve the structure of
the bond coating. Generally, an alumina scale is formed on the bond
coat by oxidation of the aluminum in the bond coat 12, or
alternatively may be formed directly on an aluminum-containing
substrate 14 without utilizing a bond coat. The thermal barrier
coating layer or undercoat 18 comprised of a ceramic material such
as YSZ, is then deposited on the bond coat 12, for example by
electron beam physical vapor deposition (EB-PVD) or other known
methods. The TBC undercoating 18 of FIGS. 1 and 2 is only
schematically represented. As known in the art, one or more of the
individual layers of the TBC undercoating 18 may have a
strain-tolerant microstructure of columnar grains as a result of
being deposited by a physical vapor deposition technique, such as
EBPVD. Alternatively, one or more of the layers may have a
noncolumnar structure as a result of being deposited by such
methods as plasma spraying, including air plasma spraying (APS).
Layers of this type are in the form of molten "splats," resulting
in a microstructure characterized by irregular flattened grains and
a degree of inhomogeneity and porosity, which features are
deleterious to thermal protection properties of the undercoating
18.
[0019] After deposition of the TBC undercoating 18 and before
application of a heat-absorbing topcoat 20, 120, the exposed
surface of the thermal barrier coating 18 can be adjusted in order
to provide a predetermined roughness of the exposed surface. Any
known methods of measuring the roughness of the exposed surface may
be utilized, and any known methods for polishing or roughening the
exposed surface may be utilized to obtain a roughness that is about
equal to the predetermined roughness. While the roughness may be
adjusted to suit the particular type of TBC material, a preferred
roughness is between about 80-125 microinches for a YSZ TBC. The
exposed TBC undercoating 18 surface may further also be cleaned to
remove debris and contaminants from the exposed surface, such as by
application of acetone or other solvents, followed by drying or
heating to remove any residual solvent before applying a
heat-absorbing topcoat layer onto the exposed surface so as to
overlie the thermal barrier coating and form a continuous overcoat
on the thermal barrier coating.
[0020] The present invention provides several different approaches
to depositing the heat-absorbing topcoat 20, 120 into the coating
system of this invention. Contrary to expected results of adding a
heat-absorbing material to a component designed to shed heat by
convective transfer, it has been unexpectedly determined that the
deposition of a heat-absorbing topcoat 20, 120 over a TBC undercoat
18 reduces the thermal transparency of the TBC undercoat 18 by
absorbing heat and preventing its direct transmission to the TBC
undercoating 18, thus subjecting the component substrate 14 to less
heat radiation than a TBC without the heat-absorbing topcoat 20,
120. These effects significantly reduce the heat load on the
substrate 14 such that an established active air cooling flow can
produce a relatively low substrate temperature in a high-heat
environment such as the flowpath of a gas turbine engine. In many
applications of interest, such as combustor liners, exhaust nozzle
liners, and other turbine engine flowpath parts, design of the
hardware and cooling flow is intended to reduce the surface
temperature (in this case of the heat-absorbing layer 20, 120)
significantly below, for example, the flame temperature in a
combustor. For example, when the topcoat 20, 120 is applied over a
TBC-coated turbine blade having cooling holes therein, the cooling
holes act in a synergistic manner to produce a cooling airflow
across the topcoat 20, 120, allowing the topcoat 20, 120 to
convectively shed absorbed heat.
[0021] The present invention provides compositions and structures
for overcoating of a substrate 14 having a TBC undercoating 18
deposited thereon. The overcoating, here a topcoat 20, 120, reduces
the thermal conductivity of the coating system 10, 110 by including
at least one heat-absorbent material. Exemplary topcoat 20, 120
component materials may comprise alumina, hafnia, tantala, silica,
platinum, nickel, iron, cobalt, chromium oxide, rare earth oxides,
silicides, carbides, and any combination thereof. The topcoat 20,
120 may further comprise a metallic-ceramic coating matrix that is
formed by the deposition, such as by spraying, of a liquid mixture
of at least one metallic element and at least one a ceramic
precursor binder, followed by heating to a degree sufficient to
thermally convert the binder and partially oxidize at least a
portion of the metallic element(s) so as to form a heat-absorbing
ceramic-metallic matrix coating 20, 120.
[0022] For example, the topcoat precursor may comprise a liquid or
semi-liquid mixture of cobalt, iron, chromium (as the metallic
element) and a ceramic precursor binder component such as silicone,
any metal organic compound that can be heated or reacted with
catalytic agents to form metal-organic type plastic materials,
including but not limited to alumoxanes such as
methoxyethoxyacetate alumoxane or acetate alumoxane, plasticized
titanium ethoxide or tantalum ethoxide, or other plasticizable
metal organic compound, wherein after applying the topcoat
precursor to the TBC undercoating 18, the topcoat 20, 120 is heat
treated (such as by heat gun or operation of the turbine engine) to
at least partially oxidize the cobalt, iron, and/or chromium to
form cobalt, iron, and/or chromium oxides and wherein the silicone
is converted by heating to form silicone dioxide in a
metallic-ceramic matrix that forms a continuous topcoat 20,
120.
[0023] Alternatively, the topcoat can be applied over a TBC in the
form of one or more thin film layers by methods such as, but not
limited to, chemical vapor deposition ("CVD"), physical vapor
deposition (PVD, plasma spray) and sputtering. Again, where a
ceramic precursor is included in the topcoat components, deposition
of the topcoat components is followed by heating to thermally
convert the ceramic precursor to yield a heat-absorbing
ceramic-metallic matrix coating.
[0024] In any embodiment, preferably, immediately upon application
and prior to thermal decomposition, the topcoat layer is comprised
of at least 25 weight percent metallic element, and at least 25
weight percent of the ceramic precursor binder component.
[0025] Exemplary embodiments of the invention are represented in
FIGS. 2 through 3. Each embodiment incorporates a topcoat layer 20,
120 over a TBC 10, 110. With reference to FIG. 2, the coating
system 10 is shown as comprising a single undercoat layer 18 lying
directly on the bond coat 12, although multiple layer TBC undercoat
layers 18 can also be provided within the scope of the invention. A
preferred composition for the TBC undercoat layer 18 is based on
binary yttria-stabilized zirconia (YSZ), a particular notable
example of which contains about 6 to about 8 weight percent yttria,
with the balance zirconia. However, other zirconia-based ceramic
materials could also be used with this invention, such as zirconia
fully stabilized by yttria, nonstabilized zirconia, or zirconia
partially or fully stabilized by ceria, magnesia, scandia and/or
other oxides. According to one aspect of the invention, a
particularly suitable material for the undercoat layer 18 is YSZ
containing about 4 to about 8 weight percent yttria (4-8% YSZ).
[0026] According to a preferred aspect of the first embodiment of
FIG. 2, the undercoat layer 18 is deposited directly over the bond
coat 12 that lies on the substrate 14. The undercoat layer 18 has
thickness that is sufficient to provide a suitable stress
distribution within the coating system 10 to promote the mechanical
integrity of the coating. A suitable thickness for this purpose is
generally on the order of about 3 to about 30 mils (about 75 to
about 750 micrometers), which is also believed to be sufficient to
provide a physical barrier to a possible reaction between the bond
coat 12 and any silica-containing layers applied over the undercoat
18. The topcoat 20 is applied over the undercoat layer 18 and is
sufficiently thick to provide the desired level of heat absorption
and thermal protection in combination with the undercoat layer 18.
While coating thickness depends on the particular application, a
suitable thickness for the topcoat 20 is also about 2 to about 30
mils (about 75 to about 750 micrometers. Preferably, the coating is
between about 8 to about 12 mils thick.
[0027] In FIG. 3, the coating system 110 differs from the
embodiment of FIG. 2 by including in the topcoat 120 a plurality of
heat-absorbing topcoat layers. As in the previous embodiment, the
TBC undercoating layer 18 lies directly on the bond coat 12. The
first topcoat layer 122 lies directly on the undercoat layer 18. A
preferred composition for the undercoating layer 18 is again based
on YSZ, preferably 3-20% YSZ. In contrast to the embodiment of FIG.
2, the topcoat 120 is formed to include at least two thin layers
122 and 124 of heat-absorbing material. Each thin topcoat layer
122, 124 may have the same composition, though it is foreseeable
that their compositions could differ. Each topcoat layer 122, 124
preferably has thicknesses of at least about 2 mils (about 50
micrometers) thickness, for a total thickness of as little 4 mils
(about 125 micrometers) and as much as 50 (about 1250 micrometers)
are foreseeable. In the embodiment of FIG. 3, the combined
thickness of the layers 122, 124 preferably constitutes at least
about one-third of the combined thickness of the coating system 10
in order for the coating system 110 to have a significant impact on
heat transparency/opacity.
[0028] Any number of TBC undercoat layers 18 and topcoat layers 20,
122 and 124 can be combined. However, the topcoat layers 20, 122
and 124 are preferably arranged so that the layer 20, 122
contacting the TBC undercoating 18 provides the highest level of
mechanical adhesion to the bond coat 12. Additionally, in the
embodiment of FIG. 3, the outer topcoat layer 124 should be
sufficiently thick to provide erosion protection to the TBC
undercoat layer 18 and any intermediate thin layer 122 or layers. A
suitable thickness for this purpose is generally on the order of up
to about 8-10 mils (about 250 micrometers).
[0029] In view of the above, it can be appreciated that each of the
TBC systems 10, 110 of this invention employs a TBC undercoat 18
whose thermal conductivity and heat transparency is reduced by the
addition of a topcoat 20, 120 having a higher heat absorption
capacity than YSZ and other conventional TBC materials.
[0030] While the invention has been described with reference to a
preferred embodiment, it will be understood by those skilled in the
art that various changes may be made and equivalents may be
substituted for elements thereof without departing from the scope
of the invention. In addition, many modifications may be made to
adapt a particular situation or material to the teachings of the
invention without departing from the essential scope thereof.
Therefore, it is intended that the invention not be limited to the
particular embodiment disclosed as the best mode contemplated for
carrying out this invention, but that the invention will include
all embodiments falling within the scope of the appended
claims.
* * * * *