U.S. patent application number 13/480906 was filed with the patent office on 2013-11-28 for turbine shroud cooling assembly for a gas turbine system.
This patent application is currently assigned to GENERAL ELECTRIC COMPANY. The applicant listed for this patent is Gregory Thomas Foster, Michelle Jessica Rogers. Invention is credited to Gregory Thomas Foster, Michelle Jessica Rogers.
Application Number | 20130315719 13/480906 |
Document ID | / |
Family ID | 48428400 |
Filed Date | 2013-11-28 |
United States Patent
Application |
20130315719 |
Kind Code |
A1 |
Rogers; Michelle Jessica ;
et al. |
November 28, 2013 |
Turbine Shroud Cooling Assembly for a Gas Turbine System
Abstract
A turbine shroud cooling assembly for a gas turbine system
includes an inner shroud component disposed within a turbine
section of the gas turbine system and proximate a hot gas path
therein, wherein the inner shroud component includes a base portion
in direct contact with the hot gas path. Also includes is a rib
protruding radially away from the base portion and disposed
proximate at least one cavity configured to receive a cooling flow
from a cooling source, wherein the cooling flow passes through the
main passage of the rib for cooling the inner shroud component.
Inventors: |
Rogers; Michelle Jessica;
(Greenville, SC) ; Foster; Gregory Thomas; (Greer,
SC) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Rogers; Michelle Jessica
Foster; Gregory Thomas |
Greenville
Greer |
SC
SC |
US
US |
|
|
Assignee: |
GENERAL ELECTRIC COMPANY
Schenectady
NY
|
Family ID: |
48428400 |
Appl. No.: |
13/480906 |
Filed: |
May 25, 2012 |
Current U.S.
Class: |
415/175 |
Current CPC
Class: |
F05D 2260/22141
20130101; F01D 9/04 20130101; F05D 2240/11 20130101; F01D 11/24
20130101 |
Class at
Publication: |
415/175 |
International
Class: |
F01D 25/00 20060101
F01D025/00 |
Claims
1. A turbine shroud cooling assembly for a gas turbine system
comprising: an inner shroud component disposed within a turbine
section of the gas turbine system and proximate a hot gas path
therein, wherein the inner shroud component includes a base portion
in direct contact with the hot gas path; and at least one rib
protruding radially away from the base portion and disposed
proximate at least one cavity configured to receive a cooling flow
from a cooling source, wherein the cooling flow passes through the
main passage of the at least one rib for cooling the inner shroud
component.
2. The turbine shroud cooling assembly of claim 1, further
comprising a first side portion passage disposed within a first
side portion and a second side portion passage disposed within a
second side portion.
3. The turbine shroud cooling assembly of claim 2, wherein the main
passage extends from the first side portion passage to the second
side portion passage, wherein the cooling flow is transferred
between the first side portion passage and the second side portion
passage by passing through the main passage.
4. The turbine shroud cooling assembly of claim 2, further
comprising a leading edge of the inner shroud component and a
trailing edge of the inner shroud component.
5. The turbine shroud cooling assembly of claim 2, further
comprising at least one fore passage disposed proximate a leading
edge and at least one aft passage disposed proximate a trailing
edge, wherein the main passage extends from the at least one fore
passage to the at least one aft passage.
6. The turbine shroud cooling assembly of claim 5, wherein the at
least one fore passage extends from the first side portion passage
to the second side portion passage, wherein the at least one aft
passage extends from the first side portion passage to the second
side portion passage.
7. The turbine shroud cooling assembly of claim 5, wherein the
cooling flow is free to transfer between the first side portion
passage, the second side portion passage, the at least one fore
passage, the at least one aft passage, and the main passage in a
continuous interconnected cooling flow circuit.
8. The turbine shroud cooling assembly of claim 1, further
comprising a plurality of channels extending from the at least one
cavity to the main passage to direct the cooling flow into the main
passage.
9. The turbine shroud cooling assembly of claim 1, wherein the main
passage includes at least one break forming a plurality of main
passages.
10. A turbine shroud cooling assembly for a gas turbine system
comprising: an inner shroud component disposed within a turbine
section of the gas turbine system and proximate a hot gas path
therein, wherein the inner shroud component includes a leading edge
and a trailing edge disposed at an aft location of the inner shroud
component relative to the leading edge; a base portion extending
from the leading edge to the trailing edge, wherein the base
portion is in direct contact with the hot gas path; and a rib
extending from a first side portion to a second side portion and
radially outward from the base portion, wherein the rib includes a
main passage extending between the first side portion and the
second side portion and configured to receive a cooling flow from a
cooling source.
11. The turbine shroud cooling assembly of claim 10, wherein the
rib is at least partially surrounded by at least one cavity
configured to receive the cooling flow from the cooling source.
12. The turbine shroud cooling assembly of claim 11, further
comprising a plurality of channels extending from the at least one
cavity to the main passage to direct the cooling flow into the main
passage.
13. The turbine shroud cooling assembly of claim 12, wherein the
plurality of channels are drilled through holes.
14. The turbine shroud cooling assembly of claim 12, further
comprising a first side portion passage disposed within the first
side portion and a second side portion passage disposed within the
second side portion, wherein the main passage extends from the
first side portion passage to the second side portion passage,
wherein the cooling flow is transferred between the first side
portion passage and the second side portion passage by passing
through the main passage.
15. The turbine shroud cooling assembly of claim 14, further
comprising a fore passage disposed proximate the leading edge and
an aft passage disposed proximate the trailing edge.
16. The turbine shroud cooling assembly of claim 15, wherein the
fore passage extends from the first side portion passage to the
second side portion passage, wherein the aft passage extends from
the first side portion passage to the second side portion
passage.
17. The turbine shroud cooling assembly of claim 15, wherein the
cooling flow is free to transfer between the first side portion
passage, the second side portion passage, the fore passage, the aft
passage, and the main passage in a continuous interconnected
cooling flow circuit.
18. A gas turbine system comprising: a compressor for distributing
a cooling flow at a high pressure; a turbine casing operably
supporting a turbine shroud assembly for receiving the cooling flow
for cooling therein; an inner shroud component comprising a leading
edge, a trailing edge spaced axially rearward of the leading edge,
and a base portion connecting the leading edge to the trailing
edge; and a rib disposed between the leading edge and the trailing
edge, and extending between a first side portion and a second side
portion, wherein the rib includes a main passage configured to
receive the cooling flow for cooling the inner shroud
component.
19. The gas turbine system of claim 18, further comprising a first
side portion passage disposed within the first side portion and a
second side portion passage disposed within the second side
portion, wherein the main passage extends from the first side
portion passage to the second side portion passage, wherein the
cooling flow is transferred between the first side portion passage
and the second side portion passage by passing through the main
passage.
20. The gas turbine system of claim 18, wherein the rib is at least
partially surrounded by an impingement cavity configured to receive
the cooling flow, and further comprising a plurality of channels
extending from the impingement cavity to the main passage to direct
the cooling flow into the main passage.
Description
BACKGROUND OF THE INVENTION
[0001] The subject matter disclosed herein relates to gas turbine
systems, and more particularly to a turbine shroud cooling assembly
for cooling turbine shrouds of such gas turbine systems.
[0002] In gas turbine systems, a combustor converts the chemical
energy of a fuel or an air-fuel mixture into thermal energy. The
thermal energy is conveyed by a fluid, often compressed air from a
compressor, to a turbine where the thermal energy is converted to
mechanical energy. As part of the conversion process, hot gas is
flowed over and through portions of the turbine as a hot gas path.
High temperatures along the hot gas path can heat turbine
components, causing degradation of components.
[0003] A turbine shroud assembly is an example of a component that
is subjected to the hot gas path and often comprises two separate
pieces, such as an inner shroud and an outer shroud. Based on the
immediate proximity of the inner shroud to the hot gas path,
various cooling schemes have been employed to maintain the
structural integrity, as well as the intended functionality, of the
inner shroud. Such cooling schemes typically result in excessive
cooling flow from a cooling source, thereby sacrificing overall
efficiency of the gas turbine system.
BRIEF DESCRIPTION OF THE INVENTION
[0004] According to one aspect of the invention, a turbine shroud
cooling assembly for a gas turbine system includes an inner shroud
component disposed within a turbine section of the gas turbine
system and proximate a hot gas path therein, wherein the inner
shroud component includes a base portion in direct contact with the
hot gas path. Also includes is a rib protruding radially away from
the base portion and disposed proximate at least one cavity
configured to receive a cooling flow from a cooling source, wherein
the cooling flow passes through the main passage of the rib for
cooling the inner shroud component.
[0005] According to another aspect of the invention, a turbine
shroud cooling assembly for a gas turbine system includes an inner
shroud component disposed within a turbine section of the gas
turbine system and proximate a hot gas path therein, wherein the
inner shroud component includes a leading edge and a trailing edge
disposed at an aft location of the inner shroud component relative
to the leading edge. Also included is a base portion extending from
the leading edge to the trailing edge, wherein the base portion is
in direct contact with the hot gas path. Further included is a rib
extending from a first side portion to a second side portion and
radially outward from the base portion, wherein the rib includes a
main passage extending between the first side portion and the
second side portion and configured to receive a cooling flow from a
cooling source.
[0006] According to yet another aspect of the invention, a gas
turbine system includes a compressor for distributing a cooling
flow at a high pressure. Also included is a turbine casing operably
supporting a turbine shroud assembly for receiving the cooling flow
for cooling therein. Further included is an inner shroud component
comprising a leading edge, a trailing edge spaced axially rearward
of the leading edge, and a base portion connecting the leading edge
to the trailing edge. Yet further included is a rib disposed
between the leading edge and the trailing edge, and extending
between a first side portion and a second side portion, wherein the
rib includes a main passage configured to receive the cooling flow
for cooling the inner shroud component.
[0007] These and other advantages and features will become more
apparent from the following description taken in conjunction with
the drawings.
BRIEF DESCRIPTION OF THE DRAWING
[0008] The subject matter, which is regarded as the invention, is
particularly pointed out and distinctly claimed in the claims at
the conclusion of the specification. The foregoing and other
features and advantages of the invention are apparent from the
following detailed description taken in conjunction with the
accompanying drawings in which:
[0009] FIG. 1 is a schematic illustration of a gas turbine
system;
[0010] FIG. 2 is a top perspective view of an inner shroud
component of a turbine shroud assembly;
[0011] FIG. 3 is a side, cross-sectional view of the inner shroud
component having a passage extending through a rib; and
[0012] FIG. 4 is a top plan view of the inner shroud component.
[0013] The detailed description explains embodiments of the
invention, together with advantages and features, by way of example
with reference to the drawings.
DETAILED DESCRIPTION OF THE INVENTION
[0014] Referring to FIG. 1, a gas turbine system is schematically
illustrated with reference numeral 10. The gas turbine system 10
includes a compressor 12, a combustor 14, a turbine 16, a shaft 18
and a fuel nozzle 20. It is to be appreciated that one embodiment
of the gas turbine system 10 may include a plurality of compressors
12, combustors 14, turbines 16, shafts 18 and fuel nozzles 20. The
compressor 12 and the turbine 16 are coupled by the shaft 18. The
shaft 18 may be a single shaft or a plurality of shaft segments
coupled together to form the shaft 18.
[0015] The combustor 14 uses a combustible liquid and/or gas fuel,
such as natural gas or a hydrogen rich synthetic gas, to run the
gas turbine system 10. For example, fuel nozzles 20 are in fluid
communication with an air supply and a fuel supply 22. The fuel
nozzles 20 create an air-fuel mixture, and discharge the air-fuel
mixture into the combustor 14, thereby causing a combustion that
creates a hot pressurized exhaust gas. The combustor 14 directs the
hot pressurized gas through a transition piece into a turbine
nozzle (or "stage one nozzle"), and other stages of buckets and
nozzles causing rotation of the turbine 16 within a turbine casing
24. Rotation of the turbine 16 causes the shaft 18 to rotate,
thereby compressing the air as it flows into the compressor 12. In
an embodiment, hot gas path components are located in the turbine
16, where hot gas flow across the components causes creep,
oxidation, wear and thermal fatigue of turbine components.
Controlling the temperature of the hot gas path components can
reduce distress modes in the components and the efficiency of the
gas turbine system 10 increases with an increase in firing
temperature. As the firing temperature increases, the hot gas path
components need to be properly cooled to meet service life and to
effectively perform intended functionality.
[0016] Referring to FIGS. 2-4, a turbine shroud cooling assembly 30
is shown. A shroud assembly is an example of a component disposed
in the turbine 16 proximate the turbine casing 24 and subjected to
the hot gas path described in detail above, the hot gas path
referred to with numeral 32. The turbine shroud cooling assembly 30
includes an inner shroud component 34 with an inner surface 36
proximate the hot gas path 32 within the turbine 16. The turbine
shroud cooling assembly 30 also includes an outer shroud component
(not illustrated) that is generally proximate to a relatively cool
fluid and/or air in the turbine 16, with the inner shroud component
34 being operably coupled to the outer shroud component. To improve
cooling of the overall turbine shroud cooling assembly 30, a
cooling flow 38 supplied by a cooling source is introduced into the
outer shroud component and directed toward the inner shroud
component 34. Specifically, a plenum within the outer shroud
component may be present to ingest and direct the cooling flow 38
toward the inner shroud component 34.
[0017] The inner shroud component 34 includes a base portion 40
having an outer surface 42, as well as the inner surface 36 that is
directly exposed to the hot gas path 32, as described above. The
base portion 40 typically arcuately extends between a leading edge
44 and a trailing edge 46 of the inner shroud component 34. Both
the leading edge 44 and the trailing edge 46 include at least one
fastening device 48, such as a rail or clip for example, that
operably couples the inner shroud component 34 with the outer
shroud component. The inner shroud component 34 also includes a
first side portion 50 and a second side portion 52 extending along
the base portion 40 between, and connected to, the leading edge 44
and the trailing edge 46. The outer surface 42 of the base portion
40 combines with the outer shroud component to form at least one
cavity 54, such as an impingement cavity, into which the cooling
flow 38 is directed toward and into.
[0018] Numerous internal passages are formed within the inner
shroud component 34 for allowing the cooling flow 38 to pass
therethrough. A first side portion passage 60 is disposed proximate
the first side portion 50 and a second side portion passage 62 is
disposed proximate the second side portion 52. Additionally, a fore
passage 64 and an aft passage 68 may be included at locations
proximate the leading edge 44 and the trailing edge 46,
respectively. Numerous other internal passages may be provided in
addition to, or alternatively to, the internal passages described
above. In the illustrated embodiment, the first side portion
passage 60, the second side portion passage 62, the fore passage 64
and the aft passage 68 are disposed proximate the perimeter of the
inner shroud component 34.
[0019] A rib 70 integrally formed with the base portion 40
protrudes radially away from the remainder of the outer surface 42
of the base portion 40 and extends between the first side portion
50 and the second side portion 52. It is to be appreciated that in
other embodiments, the rib 70 may extend at various angles across
the base portion 40, including relatively perpendicular to that
illustrated, where the rib 70 extends from proximate the leading
edge 44 to the trailing edge 46. Irrespective of the precise
location and orientation of the rib 70, in order to effectively and
efficiently cool portions of the inner shroud component 34 other
than those proximate the perimeter, a main passage 72 is formed
within the rib 70. In the illustrated embodiment, the main passage
72 extends between, and connects with, the first side portion
passage 60 and the second side portion passage 62, thereby allowing
the cooling flow 38 to be transferred through the main passage 72,
the first side portion passage 60 and the second side portion
passage 62, in any direction. Additionally, the fore passage 64 and
the aft passage 68 extend between, and connect to, the first side
portion passage 60 and the second side portion passage 62, thereby
forming a continuous, interconnected cooling flow circuit 74. It is
to be appreciated that a discontinuous circuit may be formed by
including one or more breaks in any of the passages, including the
main passage 72, the first side portion passage 60, the second side
portion passage 62, the fore passage 64 and/or the aft passage
68.
[0020] Cooling of the inner shroud component 34 is achieved by
ingesting an airstream of the cooling flow 38 from a cooling source
(not illustrated) that provides the cooling flow 38, which may
include air, a water solution and/or a gas. The cooling flow 38 is
any suitable fluid that cools the inner shroud component 34. For
example, the cooling source is a supply of compressed air from the
compressor 12, where the compressed air is diverted from the air
supply that is routed to the combustor 14. Thus, the supply of
compressed air bypasses the combustor 14 and is used to cool the
turbine shroud cooling assembly 30. The inner shroud component 34
receives the cooling flow 38 at the at least one cavity 54 and
introduces the cooling flow 38 into at least one of the first side
portion passage 60, the second side portion passage 62, the fore
passage 64 and the aft passage 68. Such an arrangement allows the
cooling flow 38 to be transferred to the main passage 72 for
cooling therein. Furthermore, the main passage 72 may be the sole,
or an additional, ingestion point for the cooling flow 38 into the
internal passages. For example, the main passage 72 may include at
least one, but typically a plurality of channels 76 formed in the
rib 70 to fluidly connect the at least one cavity 54 and the main
passage 72. The plurality of channels 76 may be drilled or formed
in any suitable manner. One or more exit paths for the cooling flow
38 may be formed throughout one or more portions of the inner
shroud component 34 to allow dumping of the cooling flow 38 to
external regions, such as the hot gas path 32. One contemplated
location of the exit paths is through the inner surface 36 of the
inner shroud component 34.
[0021] Accordingly, the main passage 72 within the rib 70 allows
the cooling flow 38 to flow through the rib 70 that is disposed
away from the perimeter of the inner shroud component 34, thereby
leading to improved cooling of the overall inner shroud component
34. Such a feature ultimately decreases the high temperatures of
various regions of the inner shroud component 34, including an aft
edge of the rib 70. Overall gas turbine system 10 efficiency is
improved based on the reduction of the cooling flow 38 that is
required to effectively cool the inner shroud component 34.
Additionally, service life of the inner shroud component 34 is
increased due to the lower temperature experienced during exposure
to the hot gas path 32.
[0022] While the invention has been described in detail in
connection with only a limited number of embodiments, it should be
readily understood that the invention is not limited to such
disclosed embodiments. Rather, the invention can be modified to
incorporate any number of variations, alterations, substitutions or
equivalent arrangements not heretofore described, but which are
commensurate with the spirit and scope of the invention.
Additionally, while various embodiments of the invention have been
described, it is to be understood that aspects of the invention may
include only some of the described embodiments. Accordingly, the
invention is not to be seen as limited by the foregoing
description, but is only limited by the scope of the appended
claims.
* * * * *