U.S. patent application number 13/653633 was filed with the patent office on 2014-07-17 for gas turbine engine component platform cooling.
This patent application is currently assigned to UNITED TECHNOLOGIES CORPORATION. The applicant listed for this patent is United Technologies Corporation. Invention is credited to Matthew A. Devore, Eleanor D. Kaufman.
Application Number | 20140196433 13/653633 |
Document ID | / |
Family ID | 51022181 |
Filed Date | 2014-07-17 |
United States Patent
Application |
20140196433 |
Kind Code |
A1 |
Kaufman; Eleanor D. ; et
al. |
July 17, 2014 |
GAS TURBINE ENGINE COMPONENT PLATFORM COOLING
Abstract
A component for a gas turbine engine according to an exemplary
aspect of the present disclosure includes, among other things, a
platform having an outer surface and an inner surface that axially
extend between a leading edge portion and a trailing edge portion.
At least one augmentation feature is disposed on at least one of
the leading edge portion and the trailing edge portion of the outer
surface of the platform.
Inventors: |
Kaufman; Eleanor D.;
(Cromwell, CT) ; Devore; Matthew A.; (Cromwell,
CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation; |
|
|
US |
|
|
Assignee: |
UNITED TECHNOLOGIES
CORPORATION
Hartford
CT
|
Family ID: |
51022181 |
Appl. No.: |
13/653633 |
Filed: |
October 17, 2012 |
Current U.S.
Class: |
60/39.83 ; 415/1;
415/176 |
Current CPC
Class: |
F01D 25/12 20130101;
F01D 5/187 20130101; F02C 7/18 20130101; F05D 2260/2212 20130101;
F05D 2260/2214 20130101; F01D 5/145 20130101; F05D 2240/81
20130101; F05D 2260/22141 20130101; F01D 9/041 20130101; F05D
2220/32 20130101 |
Class at
Publication: |
60/39.83 ;
415/176; 415/1 |
International
Class: |
F01D 5/08 20060101
F01D005/08; F01D 25/12 20060101 F01D025/12 |
Goverment Interests
[0001] This invention was made with government support under
Contract No. F33615-03-D-2354-0017 awarded by the United States Air
Force and Contract No. N00421-99-C-1270-0011 awarded by the United
States Navy.
Claims
1. A component for a gas turbine engine, comprising: a platform
having an outer surface and an inner surface that axially extend
between a leading edge portion and a trailing edge portion; and at
least one augmentation feature disposed on at least one of said
leading edge portion and said trailing edge portion of said outer
surface of said platform.
2. The component as recited in claim 1, wherein said platform is an
inner diameter platform.
3. The component as recited in claim 1, wherein the component is a
turbine vane.
4. The component as recited in claim 1, wherein said outer surface
is a non-gas path side of said platform and said inner surface is a
gas path side of said platform.
5. The component as recited in claim 1, comprising an airfoil that
extends from said inner surface of said platform.
6. The component as recited in claim 1, wherein said at least one
augmentation feature includes a plurality of trip strips.
7. The component as recited in claim 6, wherein said plurality of
trip strips are disposed at said trailing edge portion of said
platform.
8. The component as recited in claim 6, wherein each of said
plurality of trip strips are angled relative to opposing mate faces
of said platform.
9. The component as recited in claim 6, wherein each of said
plurality of trip strips include a first portion and a second
portion that is transverse to the first portion.
10. The component as recited in claim 9, wherein said first
portions are angled at a first angle relative to a mate face of
said platform and said second portions are angled at a second angle
different from said first angle relative to said mate face.
11. The component as recited in claim 1, wherein said at least one
augmentation feature is disposed at said leading edge portion of
said platform.
12. The component as recited in claim 1, wherein said at least one
augmentation feature is disposed on a portion of said outer surface
of said platform that axially overlaps a neighboring component of
the gas turbine engine.
13. A gas turbine engine, comprising: a compressor section; a
combustor section in fluid communication with said compressor
section; a turbine section in fluid communication with said
combustor section; and wherein at least one of said compressor
section and said turbine section includes: a first component having
a platform that includes an outer surface and an inner surface that
axially extend between a leading edge portion and a trailing edge
portion; a second component mounted adjacent to said first
component and including a platform; wherein a portion of said
platform of said first component axially overlaps a portion of said
platform of said second component, and said portion of said first
platform includes at least one augmentation feature disposed on
said outer surface of said platform.
14. The gas turbine engine as recited in claim 13, wherein said
first component is a vane and said second component is a blade.
15. The gas turbine engine as recited in claim 13, wherein said at
least one augmentation feature includes a plurality of trip
strips.
16. The gas turbine engine as recited in claim 13, wherein said at
least one augmentation feature is disposed on at least one of said
leading edge portion and said trailing edge portion of said
platform of said first component.
17. The gas turbine engine as recited in claim 13, wherein said at
least one augmentation feature is angled relative to a mate face of
said platform.
18. A method of cooling a component of a gas turbine engine,
comprising the steps of: cooling a platform of the component with a
leakage airflow that is communicated from a cavity positioned
radially inwardly from the component by circulating the leakage
airflow over at least one augmentation feature that is disposed on
an outer surface of the platform.
19. The method as recited in claim 18, wherein the leakage airflow
is not a dedicated cooling airflow that is communicated inside of
the component.
20. The method as recited in claim 18, wherein the at least one
augmentation feature is disposed on a trailing edge portion of the
platform.
Description
BACKGROUND
[0002] This disclosure relates generally to a gas turbine engine,
and more particularly to a component that can be incorporated into
a gas turbine engine. The component can include platform cooling
augmentation features for cooling the platform of the
component.
[0003] Gas turbine engines typically include a compressor section,
a combustor section and a turbine section. During operation, air is
pressurized in the compressor section and is mixed with fuel and
burned in the combustor section to generate hot combustion gases.
The hot combustion gases are communicated through the turbine
section, which extracts energy from the hot combustion gases to
power the compressor section and other gas turbine engine
loads.
[0004] Both the compressor and turbine sections of the gas turbine
engine may include alternating rows of rotating blades and
stationary vanes that extend into the core flow path of the gas
turbine engine. For example, in the turbine section, turbine blades
rotate to extract energy from the hot combustion gases that are
communicated along the core flow path of the gas turbine engine.
The turbine vanes prepare the airflow for the next set of
blades.
[0005] Blades and vanes are examples of components that may need
cooled in order to withstand the relatively high temperature of the
hot combustion gases that are communicated along the core flow
path. Typically, cooling is achieved by communicating a dedicated
cooling airflow to select portions of the components.
SUMMARY
[0006] A component for a gas turbine engine according to an
exemplary aspect of the present disclosure includes, among other
things, a platform having an outer surface and an inner surface
that axially extend between a leading edge portion and a trailing
edge portion. At least one augmentation feature is disposed on at
least one of the leading edge portion and the trailing edge portion
of the outer surface of the platform.
[0007] In a further non-limiting embodiment of the foregoing gas
turbine engine, the platform is an inner diameter platform.
[0008] In a further non-limiting embodiment of either of the
foregoing gas turbine engines, the component is a turbine vane.
[0009] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, the outer surface is a non-gas path side of
the platform and the inner surface is a gas path side of the
platform.
[0010] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, an airfoil extends from the inner surface of
the platform.
[0011] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, the at least one augmentation feature includes
a plurality of trip strips.
[0012] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, the plurality of trip strips are disposed at
the trailing edge portion of the platform.
[0013] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, each of the plurality of trip strips are
angled relative to opposing mate faces of the platform.
[0014] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, each of the plurality of trip strips include a
first portion and a second portion that is transverse to the first
portion.
[0015] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, the first portions are angled at a first angle
relative to a mate face of the platform and the second portions are
angled at a second angle different from the first angle relative to
the mate face.
[0016] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, the at least one augmentation feature is
disposed at the leading edge portion of the platform.
[0017] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, the at least one augmentation feature is
disposed on a portion of the outer surface of the platform that
axially overlaps a neighboring component of the gas turbine
engine.
[0018] A gas turbine engine, according to an exemplary aspect of
the present disclosure includes, among other things, a compressor
section, a combustor section in fluid communication with the
compressor section, and a turbine section in fluid communication
with the combustor section. At least one of the compressor section
and the turbine section includes a first component having a
platform that includes an outer surface and an inner surface that
axially extend between a leading edge portion and a trailing edge
portion and a second component mounted adjacent to the first
component and including a platform. A portion of the platform of
the first component axially overlaps a portion of the platform of
the second component, and the portion of the first platform
includes at least one augmentation feature disposed on the outer
surface of the platform.
[0019] In a further non-limiting embodiment of the foregoing gas
turbine engine, the first component is a vane and the second
component is a blade.
[0020] In a further non-limiting embodiment of either of the
foregoing gas turbine engines, the at least one augmentation
feature includes a plurality of trip strips.
[0021] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, the at least one augmentation feature is
disposed on at least one of the leading edge portion and the
trailing edge portion of the platform of the first component.
[0022] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, the at least one augmentation feature is
angled relative to a mate face of the platform.
[0023] A method of cooling a component of a gas turbine engine
according to another exemplary aspect of the present disclosure
includes, among other things, cooling a platform of the component
with a leakage airflow that is communicated from a cavity
positioned radially inwardly from the component by circulating the
leakage airflow over at least one augmentation feature that is
disposed on an outer surface of the platform.
[0024] In a further non-limiting embodiment of the foregoing method
of cooling a component of a gas turbine engine, the leakage airflow
is not a dedicated cooling airflow that is communicated inside of
the component.
[0025] In a further non-limiting embodiment of either of the
foregoing methods of cooling a component of a gas turbine engine,
the at least one augmentation feature is disposed on a trailing
edge portion of the platform.
[0026] The various features and advantages of this disclosure will
become apparent to those skilled in the art from the following
detailed description. The drawings that accompany the detailed
description can be briefly described as follows.
BRIEF DESCRIPTION OF THE DRAWINGS
[0027] FIG. 1 illustrates a schematic, cross-sectional view of a
gas turbine engine.
[0028] FIG. 2 illustrates a cross-section of a portion of a gas
turbine engine.
[0029] FIG. 3 illustrates a component that can be incorporated into
a gas turbine engine.
[0030] FIG. 4 illustrates another view of the component of FIG.
3.
[0031] FIG. 5 illustrates a platform of a component.
[0032] FIG. 6 illustrates another platform of a component.
[0033] FIG. 7 illustrates augmentation features that can be
incorporated into a component of a gas turbine engine.
[0034] FIG. 8 illustrates additional augmentation features that can
be incorporated into a component of a gas turbine engine.
DETAILED DESCRIPTION
[0035] FIG. 1 schematically illustrates a gas turbine engine 20.
The exemplary gas turbine engine 20 is a two-spool turbofan engine
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 augmenter section (not shown) among other
systems for features. The fan section 22 drives air along a bypass
flow path B, while the compressor section 24 drives air along a
core flow path C for compression and communication into the
combustor section 26. The hot combustion gases generated in the
combustor section 26 are expanded 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 turbofan engines and these
teachings could extend to other types of engines, including but not
limited to, three-spool engine architectures.
[0036] The gas turbine engine 20 generally includes a low speed
spool 30 and a high speed spool 32 mounted for rotation about an
engine centerline longitudinal axis A. The low speed spool 30 and
the high speed spool 32 may be mounted relative to an engine static
structure 33 via several bearing systems 31. It should be
understood that other bearing systems 31 may alternatively or
additionally be provided.
[0037] The low speed spool 30 generally includes an inner shaft 34
that interconnects a fan 36, a low pressure compressor 38 and a low
pressure turbine 39. The inner shaft 34 can be connected to the fan
36 through a geared architecture 45 to drive the fan 36 at a lower
speed than the low speed spool 30. The high speed spool 32 includes
an outer shaft 35 that interconnects a high pressure compressor 37
and a high pressure turbine 40. In this embodiment, the inner shaft
34 and the outer shaft 35 are supported at various axial locations
by bearing systems 31 positioned within the engine static structure
33.
[0038] A combustor 42 is arranged between the high pressure
compressor 37 and the high pressure turbine 40. A mid-turbine frame
44 may be arranged generally between the high pressure turbine 40
and the low pressure turbine 39. The mid-turbine frame 44 can
support one or more bearing systems 31 of the turbine section 28.
The mid-turbine frame 44 may include one or more airfoils 46 that
extend within the core flow path C.
[0039] The inner shaft 34 and the outer shaft 35 are concentric and
rotate via the bearing systems 31 about the engine centerline
longitudinal axis A, which is co-linear with their longitudinal
axes. The core airflow is compressed by the low pressure compressor
38 and the high pressure compressor 37, is mixed with fuel and
burned in the combustor 42, and is then expanded over the high
pressure turbine 40 and the low pressure turbine 39. The high
pressure turbine 40 and the low pressure turbine 39 rotationally
drive the respective high speed spool 32 and the low speed spool 30
in response to the expansion.
[0040] Each of the compressor section 24 and the turbine section 28
may include alternating rows of rotor assemblies and vane
assemblies (shown schematically) that carry airfoils that extend
into the core flow path C. For example, the rotor assemblies can
carry a plurality of rotating blades 25, while each vane assembly
can carry a plurality of vanes 27 that extend into the core flow
path C. The blades 25 of the rotor assemblies create or extract
energy (in the form of pressure) from the core airflow that is
communicated through the gas turbine engine 20 along the core flow
path C. The vanes 27 of the vane assemblies direct the core air
flow to the blades 25 to either add or extract energy.
[0041] Various components of the gas turbine engine 20, such as the
blades 25 and the vanes 27 on the compressor section 24 and/or the
turbine section 28, may be subjected to repetitive thermal cycling
under widely ranging temperatures and pressures. The hardware of
the turbine section 28 is particularly subjected to relatively
extreme operating conditions. Therefore, some components may need
cooled during engine operation. Example cooling features that can
be incorporated into the components to improve cooling efficiency
are described below.
[0042] FIG. 2 schematically illustrates a portion 100 of a gas
turbine engine, such as the gas turbine engine 20 of FIG. 1. In
this exemplary embodiment, the portion 100 represents part of the
turbine section 28 of the gas turbine engine 20. However, it should
be understood that other portions of the gas turbine engine 20
could benefit from the teachings of this disclosure, including but
not limited to, the compressor section 24.
[0043] In this exemplary embodiment, the portion 100 includes a
first component 50 and a second component 52 positioned adjacent to
the first component 50. For example, the first component 50 may
represent a vane (a generally static structure) and the second
component 52 may represent a blade mounted for rotation about the
engine centerline longitudinal axis A. Although only a single vane
and a single blade are illustrated in FIG. 2, multiple vanes and
blades could be circumferentially disposed about the engine
centerline longitudinal axis A to provide vane and rotor
assemblies. The portion 100 could also include additional,
alternating rows of vanes and blades. FIG. 2 is highly schematic,
and it should be understood that the various features depicted by
this figure are not necessarily drawn to the scale they would be in
practice.
[0044] In this embodiment, the first component 50 establishes a
radially outer and radially inner flow path boundary of the core
flow path C and directs the hot combustion gases communicated along
the core flow path C to the second component 52. The second
component 52 rotates to extract energy from the hot combustion
gases that are communicated through the gas turbine engine 20.
[0045] The first component 50 and the second component 52 are
mounted within the portion 100 such that a gap 56 extends between
the first component 50 and the second component 52. A positive
pressure can be maintained within the portion 100 by communicating
a leakage airflow 60 into the gap 56. The leakage airflow 60 is
communicated through a cavity 58 (positioned radially inwardly from
the first component 50 and the second component 52) and then
through the gap 56 to keep the hot combustion gases of the core
flow path C from entering through the gap 56 and potentially
damaging components. The leakage airflow 60 may be communicated
from the compressor section 24 or some other upstream location of
the gas turbine engine 20.
[0046] As is discussed in greater detail below, the leakage airflow
60 can also be used to cool portions of one or both of the first
component 50 and the second component 52. In other words, the
leakage airflow 60 that may be used to cool portions of the first
component 50 and/or the second component 52 is not a dedicated
cooling airflow that serves no other purpose other than to cool the
component(s) 50, 52. In this disclosure, the term "dedicated
cooling airflow" may refer to air which feeds the inside of the
component(s) 50, 52 and the term "leakage airflow" may refer to
airflow that bypasses the inside of the component(s) 50, 52, such
as for purging cavities or preventing ingestion.
[0047] The first component 50 and/or the second component 52 can
include cooling features that increase a local heat transfer effect
of the first and/or second component 50, 52 without requiring a
large flow pressure ratio. For example, in one embodiment, the
first component 50 includes a platform 64A and the second component
52 includes a platform 64B. Each of the platforms 64A, 64B includes
an outer surface 74 and an inner surface 76. In one embodiment, at
least a portion of the second component 52 extends radially inward
from, or under, the first component 50.
[0048] The platform 64A of the first component may include one or
more augmentation features 78 disposed on the outer surface 74 for
increasing the heat transfer effect between the platform 64A and
the leakage airflow 60. The platform 64A can be cooled by
circulating the leakage airflow 60 over the augmentation features
78. In one embodiment, the augmentation features 78 are disposed on
a portion 55 of the platform 64A of the first component 50 that
axially overlaps the platform 64B of the second component 52. The
platform 64B could also include one or more augmentation features
on its outer surface 74, although not shown in this embodiment.
[0049] FIGS. 3 and 4 illustrate a component 150 that can be
incorporated into a gas turbine engine, such as the gas turbine
engine 20 of FIG. 1. In this embodiment, the component 150 is a
turbine vane similar to the first component 50 of FIG. 2. However,
the various features described herein with respect to the component
150 could extend to other components of the gas turbine engine 20,
including but not limited to blades (i.e., the second component 52
of FIG. 2).
[0050] The component 150 of this embodiment includes an inner
diameter platform 62, an outer diameter platform 64 and an airfoil
66 that extends between the inner diameter platform 62 and the
outer diameter platform 64. Each of the inner diameter platform 62
and the outer diameter platform 64 includes a leading edge portion
68, a trailing edge portion 70 and opposing mate faces 71, 73. The
inner diameter and outer diameter platforms 62, 64 axially extend
between the leading edge portion 68 and the trailing edge portion
70 and circumferentially extend between the opposing mate faces 71,
73. The opposing mate faces 71, 73 can be mounted relative to
corresponding mate faces of adjacent components of a gas turbine
engine to provide a full ring assembly, such as a full ring vane
assembly that can be circumferentially disposed about the engine
centerline longitudinal axis A (see FIG. 1).
[0051] The inner diameter and outer diameter platforms 62, 64 can
also include an outer surface 74 (for example, a non-gas path side)
and an inner surface 76 (a gas path side). In other words, when the
component 150 is mounted within a gas turbine engine 20, the outer
surfaces 74 are positioned on a non-core flow path side of the
component 150, while the inner surfaces 76 establish the outer
boundaries of the core flow path C of the gas turbine engine
20.
[0052] One or both of the inner diameter platform 62 and the outer
diameter platform 64 can include one or more augmentation features
78 disposed on the outer surfaces 74 of the inner diameter platform
62 and/or the outer diameter platform 64. In this embodiment, the
augmentation features 78 are positioned at the trailing edge
portion 70 of the inner diameter platform 62 (see FIG. 2 and FIG.
4). In another embodiment, the augmentation features 78 may be
disposed at the leading edge portion 68 of the inner diameter
platform 62 and/or the outer diameter platform 64 (see FIG. 5). In
yet another embodiment, the augmentation features 78 may be formed
on both the leading edge portion 68 and the trailing edge portion
70 of the inner diameter platform 62 and/or the outer diameter
platform 64 (see FIG. 6). The augmentation features 78 may be
disposed on any portion of the inner diameter platform 62 and/or
the outer diameter platform 64, including any portion of the inner
or outer diameter platforms 62, 64 that axially overlap a
neighboring component (see portion 55 of FIG. 2, for example).
[0053] The exemplary augmentation features 78 can include any heat
transfer augmentation features including but not limited to trip
strips, pin fins, chevron trip strips, or any combination of
features. In this embodiment, the augmentation features 78 include
trip strips. The augmentation features 78 turbulate the flow of the
leakage airflow 60 that comes into contact with the component 150
(as shown in FIG. 2) and adds surface area to platform(s) 62, 64 to
enhance heat transfer between the leakage airflow 60 and the
platform(s) 62, 64.
[0054] FIG. 7 illustrates additional features of the augmentation
features 78 described above. The illustrated platform 62, 64 could
be representative of either an inner diameter platform or an outer
diameter platform. In this embodiment, the augmentation features 78
include a plurality of trip strips 78A that are circumferentially
spaced between the opposing mate faces 71, 73 at a trailing edge
portion 70 of the platform 62, 64. The trip strips 78A extend
parallel to one another and may each be angled at an angle .alpha.
relative to the opposing mate faces 71, 73. The value of the angle
.alpha. can vary depending on design specific criteria, including
but not limited to the amount of heat transfer required to cool the
platform 62, 64. The leakage airflow 60 may be circulated over the
trip strips 78A in both a circumferential direction CD as well as
an axial direction AD to cool the platform 62, 64.
[0055] FIG. 8 illustrates another plurality of augmentation
features 178 that can be incorporated into a platform 62, 64 of a
component 150. In this embodiment, the augmentation features 178
include a plurality of trip strips 178A that are circumferentially
spaced between the opposing mate faces 71, 73 at a trailing edge
portion 70 of the platform 62, 64. The plurality of trip strips
178A of this embodiment include a first portion 80 and a second
portion 82 that is transverse to the first portion 80. The first
portions 80 of the plurality of trip strips 178A extend parallel to
one another and may each be angled at an angle .alpha. relative to
the opposing mate faces 71, 73. The second portions 82 of the
plurality of trip strips 178A also extend parallel to one another
and may be angled at an angle .beta. relative to the opposing mate
faces 71, 73. In this embodiment, the angle .beta. is a different
angle than the angle .alpha.. The second portions 82 of the
plurality of trip strips 178A direct the leakage airflow 60 into
the core flow path C by directing the leakage airflow 60
circumferentially along the direction of rotation of a neighboring
component (shown schematically via arrow 110), such as a blade (see
the second component 52 of FIG. 2), thereby providing improved heat
transfer and reducing mixing loss.
[0056] Although the different non-limiting embodiments are
illustrated as having specific components, the embodiments of this
disclosure are not limited to those particular combinations. It is
possible to use some of the components or features from any of the
non-limiting embodiments in combination with features or components
from any of the non-limiting embodiments.
[0057] 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 and illustrated in these
exemplary embodiments, other arrangements could also benefit from
the teachings of this disclosure.
[0058] The foregoing description shall be interpreted as
illustrative and not in any limiting sense. A worker of ordinary
skill in the art would recognize that various modifications could
come within the scope of this disclosure. For these reasons, the
following claims should be studied to determine the true scope and
content of this disclosure.
* * * * *