U.S. patent application number 13/162009 was filed with the patent office on 2012-12-20 for cell structure thermal barrier coating.
Invention is credited to James A. Dierberger.
Application Number | 20120317984 13/162009 |
Document ID | / |
Family ID | 46207910 |
Filed Date | 2012-12-20 |
United States Patent
Application |
20120317984 |
Kind Code |
A1 |
Dierberger; James A. |
December 20, 2012 |
CELL STRUCTURE THERMAL BARRIER COATING
Abstract
A combustor component of a gas turbine engine includes a thermal
barrier coating on a substrate, the thermal barrier coating defines
a multiple of cells.
Inventors: |
Dierberger; James A.;
(Hebron, CT) |
Family ID: |
46207910 |
Appl. No.: |
13/162009 |
Filed: |
June 16, 2011 |
Current U.S.
Class: |
60/722 |
Current CPC
Class: |
Y02T 50/60 20130101;
C23C 4/18 20130101; C23C 4/01 20160101; F23M 2900/05004 20130101;
F23M 5/00 20130101; F23M 5/04 20130101; F05D 2250/283 20130101;
F23R 3/007 20130101; F23R 2900/00018 20130101; Y02T 50/6765
20180501 |
Class at
Publication: |
60/722 |
International
Class: |
F23D 99/00 20100101
F23D099/00 |
Claims
1. A component of a gas turbine engine comprising: a substrate; and
a thermal barrier coating on said substrate, said thermal barrier
coating defines a multiple of cells.
2. The component as recited in claim 1, wherein said substrate
defines an offset within each of said multiple of cells.
3. The component as recited in claim 2, wherein said offset within
each of said multiple of cells is a post.
4. The component as recited in claim 3, wherein said post within
each of said multiple of cells is frustro-conical.
5. The component as recited in claim 3, wherein said post is
located within a center of each of said multiple of cells.
6. The component as recited in claim 3, wherein said post defines a
height that is approximately two-thirds the thickness of said
thermal barrier coating.
7. The component as recited in claim 2, wherein said offset within
each of said multiple of cells is a divot.
8. The component as recited in claim 7, wherein said divot is
located within a center of each of said multiple of cells.
9. The combustor component as recited in claim 1, wherein each of
said multiple of cells are hexagonal in shape.
10. The combustor component as recited in claim 1, wherein each of
said multiple of cells are separated by a gap.
11. The combustor component as recited in claim 1, said substrate
is a nickel base superalloy.
12. A combustor component of a gas turbine engine comprising: a
substrate which defines a multiple of offsets; and a thermal
barrier coating on said substrate, said thermal barrier coating
defines a multiple of cells, each of said multiple of cells
correspond with at least one of said multiple of offsets.
13. The combustor component as recited in claim 12, wherein at
least one of said multiple of offsets is a post.
14. The combustor component as recited in claim 13, wherein said
post is frustro-conical.
15. The combustor component as recited in claim 13, wherein each of
said multiple of posts is located within a center of each of said
multiple of cells.
16. The combustor component as recited in claim 13, wherein said
post is approximately two-thirds the thickness of said thermal
barrier coating.
17. The combustor component as recited in claim 12, wherein at
least one of said multiple of offsets is a divot.
18. The combustor component as recited in claim 17, wherein s each
of said multiple of divots is located within a center of each of
said multiple of cells
19. The combustor component as recited in claim 12, wherein each of
said multiple of cells are hexagonal in shape.
20. The combustor component as recited in claim 19, wherein each of
said multiple of cells are separated by a gap.
Description
BACKGROUND
[0001] The present disclosure relates to a thermal barrier coating,
and more particularly to a combustor with a thermal barrier
coating.
[0002] A gas turbine engine includes a compressor for compressing
air which is mixed with a fuel and channeled to a combustor wherein
the mixture is ignited within a combustion chamber to generate hot
combustion core gases. At least some combustors include combustor
liners to channel the combustion gases to a turbine which extracts
energy from the combustion core gases to power the compressor, as
well as produce useful work to propel an aircraft in flight or to
power a load, such as an electrical generator.
[0003] The combustor liners often include a thermal barrier coating
to increase durability. The difference in properties between the
ceramic thermal barriers and the metal substrates to which the
ceramic is applied may lead to mismatched strains which ultimately
lead to areas of coating spallation which may tend to spall or
flake. Once spalled, substrate degradation in the form of cracking
and oxidation may follow.
SUMMARY
[0004] A component of a gas turbine engine according to an
exemplary aspect of the present disclosure includes a thermal
barrier coating on a substrate, the thermal barrier coating defines
a multiple of cells.
[0005] A combustor component of a gas turbine engine according to
an exemplary aspect of the present disclosure includes a substrate
which defines a multiple of offsets and a thermal barrier coating
on the substrate, the thermal barrier coating defines a multiple of
cells, each of the multiple of cells correspond with at least one
of the multiple of offsets.
BRIEF DESCRIPTION OF THE DRAWINGS
[0006] Various features will become apparent to those skilled in
the art from the following detailed description of the disclosed
non-limiting embodiment. The drawings that accompany the detailed
description can be briefly described as follows:
[0007] FIG. 1 is a schematic cross-section of a gas turbine
engine;
[0008] FIG. 2 is a perspective partial sectional view of an
exemplary annular combustor that may be used with the gas turbine
engine shown in FIG. 1;
[0009] FIG. 3 is a cross-sectional view of an exemplary combustor
that may be used with the gas turbine engine shown in FIG. 2;
[0010] FIG. 4 is a facial view of a combustor component;
[0011] FIG. 5 is a cross-sectional view of one non-limiting
embodiment of the combustor component; and
[0012] FIG. 6 is a cross-sectional view of another non-limiting
embodiment of the combustor component.
DETAILED DESCRIPTION
[0013] FIG. 1 schematically illustrates a gas turbine engine 20.
The gas turbine engine 20 is disclosed herein as a two-spool
turbofan that generally incorporates a fan section 22, a compressor
section 24, a combustor section 26 and a turbine section 28.
Alternative engines might include an augmentor section (not shown)
among other systems or features. The fan section 22 drives air
along a bypass flowpath while the compressor section 24 drives air
along a core flowpath for compression and communication into the
combustor section 26 then expansion through the turbine section 28.
Although depicted as a turbofan gas turbine engine in the disclosed
non-limiting embodiment, it should be understood that the concepts
described herein are not limited to use with turbofans as the
teachings may be applied to other types of turbine engines.
[0014] The engine 20 generally includes a low speed spool 30 and a
high speed spool 32 mounted for rotation about an engine central
longitudinal axis A relative to an engine static structure 36 via
several bearing systems 38. It should be understood that various
bearing systems 38 at various locations may alternatively or
additionally be provided.
[0015] The low speed spool 30 generally includes an inner shaft 40
that interconnects a fan 42, a low pressure compressor 44 and a low
pressure turbine 46. The inner shaft 40 is connected to the fan 42
through a geared architecture 48 to drive the fan 42 at a lower
speed than the low speed spool 30. The high speed spool 32 includes
an outer shaft 50 that interconnects a high pressure compressor 52
and high pressure turbine 54. A combustor 56 is arranged between
the high pressure compressor 52 and the high pressure turbine 54.
The inner shaft 40 and the outer shaft 50 are concentric and rotate
about the engine central longitudinal axis A which is collinear
with their longitudinal axes.
[0016] The core airflow is compressed by the low pressure
compressor 44 then the high pressure compressor 52, mixed and
burned with fuel in the combustor 56, then expanded over the high
pressure turbine 54 and low pressure turbine 46. The turbines 54,
46 rotationally drive the respective low speed spool 30 and high
speed spool 32 in response to the expansion.
[0017] With reference to FIG. 2, the combustor 56 generally
includes an outer liner 60 and an inner liner 62. It should be
understood that various combustor arrangements such as a can
combustor as well as other high temperature components such as
turbine components may alternatively benefit herefrom.
[0018] With reference to FIG. 3, outer liner 60 and inner liner 62
are spaced inward from a combustor case 64 such that a combustion
chamber 68 is defined between liners 60, 62. The outer liner 60 and
combustor case 64 define an outer passageway 70. The inner liner 62
and combustor case 64 define an inner passageway 72. Combustion
chamber 68 is generally annular in shape and is defined between
liners 60, 62. Outer and inner liners 60, 62 extend toward the
turbine section 28.
[0019] The outer and inner liners 60, 62 support a multiple of
liner panels 74. Each liner panel 74 generally includes a metallic
substrate 76 with a thermal barrier coating 78 (FIG. 4) on an inner
surface 80 which faces the combustion chamber 68 (FIG. 5). It
should be understood that although a particular combustor is
illustrated, other combustor types with various combustor liner
panel arrangements will also benefit herefrom. It should be further
understood that the disclosed liner panel is but a single
illustrated embodiment and should not be limited only thereto such
that the disclosed liner panel 74 may be considered but one
combustor component of various types manufactured of a substrate 76
upon which the thermal barrier coating 78 is applied as disclosed
herein.
[0020] The substrate 76 may be a nickel base superalloy, other
metallic material, or Ceramic Matrix Composite material. The
thermal barrier coating 78 may be applied in, for example, a plasma
spray coating process in which powders are injected into a high
temperature, high velocity stream of ionized gases. At the point
where the powders are injected into the gas stream, the temperature
can be about 15,000.degrees F (8315 C). As a result, the powders
are typically molten when they strike the surface of the substrate
forming an interlocking "splat" type structure. It should be
understood that the thermal barrier coating 78 may be sprayed on a
bondcoat which has been applied to the substrate 76 which has been
found to improve adhesion and is well known in the industry.
[0021] With reference to FIG. 4, the thermal barrier coating 78 is
applied to the substrate 76 (FIG. 5) in a manner to form a grid
pattern which includes a multiple of cells 82 each separated by a
narrow gap 84 (also illustrate in FIG. 5). The gap 84 may be formed
to be exceedingly narrow yet still facilities thermal barrier
coating 78 strain tolerance as the gaps 84 define the maximum size
of potential `mudflat` cracks that may occur due to sintering. This
segregation facilitates durability and accommodation of thermal
gradients.
[0022] Each cell 82 may be of a particular shape such as hexagonal
(shown), square, triangular or other shape. In one non-limiting
embodiment, each cell 82 may be of approximately 0.25 inches (6.35
mm) across opposed corners of the illustrated hexagonal shape.
[0023] The multiple of cells 82 with the respective narrow gap 84
may be manufactured, for example, with a matrix grid of a polyester
fugitive material which is applied over the substrate 76 prior to
the "splat" type plasma spray process then thereafter baked out to
form the gaps 84. The gap 84 could be as narrow as a crack and
still provide a measure of strain relief and durability improvement
such that an alternative thermal process may include a laser
testament to pre-treat the ceramic and produce the matrix grid of
gaps 84. It should be understood that various processes may
alternatively or additionally be utilized to essentially mask or
mark the substrate 76 to form the matrix grid of gaps 84.
[0024] An offset 86 interfaces with each of the multiple of cells
82. The offset 86 may extend outward from the surface 80 of the
substrate 76 to form a post which extends into the thermal barrier
coating 78 (FIG. 5). Alternatively, an offset 86' may extend into
the surface 64 of the substrate 76' to form a divot which at least
partially receives the thermal barrier coating 78 (FIG. 6). In one
disclosed non-limiting embodiment, the offset 86, 86' defines a
reverse taper such as a frustro-conical structure relative to the
surface 80 of the substrate 76 that is two-thirds (2/3) the
thickness of the thermal barrier coating 78 to form an interlock
for the thermal barrier coating 78 at each cell 82.
[0025] The offset 86, 86' provides a reference point about which
sintering shrinkage of the thermal barrier coating 78 will
interlock to facilitate adhesion of the thermal barrier coating 78
to the substrate 76. That is, the thermal barrier coating 78, when
sintered, mechanically interlocks onto the post or into the divot.
This mechanical interlock significantly increases the life of the
thermal barrier coating 78 and therefore increases the life of the
components onto which the thermal barrier coating 78 is applied. As
each cell 82 is provided with an offset 86, 86', adhesion to the
substrate 76 is supplemented and the thermal barrier coating 78 may
last for the entire life of the component which is protected
thereby.
[0026] It should be understood that relative positional terms such
as "forward," "aft," "upper," "lower," "above," "below," and the
like are with reference to the normal operational attitude of the
vehicle and should not be considered otherwise limiting.
[0027] It should be understood that like reference numerals
identify corresponding or similar elements throughout the several
drawings. It should also be understood that although a particular
component arrangement is disclosed in the illustrated embodiment,
other arrangements will benefit herefrom.
[0028] Although particular step sequences are shown, described, and
claimed, it should be understood that steps may be performed in any
order, separated or combined unless otherwise indicated and will
still benefit from the present disclosure.
[0029] The foregoing description is exemplary rather than defined
by the limitations within. Various non-limiting embodiments are
disclosed herein, however, one of ordinary skill in the art would
recognize that various modifications and variations in light of the
above teachings will fall within the scope of the appended claims.
It is therefore to be understood that within the scope of the
appended claims, the disclosure may be practiced other than as
specifically described. For that reason the appended claims should
be studied to determine true scope and content.
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