U.S. patent application number 13/205763 was filed with the patent office on 2013-02-14 for turbine airfoil and method of controlling a temperature of a turbine airfoil.
This patent application is currently assigned to GENERAL ELECTRIC COMPANY. The applicant listed for this patent is Bradley Taylor Boyer, Anthony Louis Giglio, Ross James Gustafson, Christopher Michael Penny, Aaron Ezekiel Smith. Invention is credited to Bradley Taylor Boyer, Anthony Louis Giglio, Ross James Gustafson, Christopher Michael Penny, Aaron Ezekiel Smith.
Application Number | 20130039758 13/205763 |
Document ID | / |
Family ID | 46639383 |
Filed Date | 2013-02-14 |
United States Patent
Application |
20130039758 |
Kind Code |
A1 |
Smith; Aaron Ezekiel ; et
al. |
February 14, 2013 |
TURBINE AIRFOIL AND METHOD OF CONTROLLING A TEMPERATURE OF A
TURBINE AIRFOIL
Abstract
According to one aspect of the invention, a turbine airfoil
includes a platform and a blade extending from the platform. The
airfoil also includes a slot formed in a slashface of the platform,
the slot being configured to receive a pressurized fluid via
passages and configured to direct the pressurized fluid to a
selected region of the turbine airfoil to improve airfoil life.
Inventors: |
Smith; Aaron Ezekiel;
(Simpsonville, SC) ; Boyer; Bradley Taylor;
(Greenville, SC) ; Giglio; Anthony Louis;
(Simpsonville, SC) ; Gustafson; Ross James;
(Greenville, SC) ; Penny; Christopher Michael;
(Greer, SC) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Smith; Aaron Ezekiel
Boyer; Bradley Taylor
Giglio; Anthony Louis
Gustafson; Ross James
Penny; Christopher Michael |
Simpsonville
Greenville
Simpsonville
Greenville
Greer |
SC
SC
SC
SC
SC |
US
US
US
US
US |
|
|
Assignee: |
GENERAL ELECTRIC COMPANY
Schenectady
NY
|
Family ID: |
46639383 |
Appl. No.: |
13/205763 |
Filed: |
August 9, 2011 |
Current U.S.
Class: |
416/1 ;
416/97R |
Current CPC
Class: |
F05D 2260/205 20130101;
F05D 2240/81 20130101; F05D 2260/2212 20130101; F01D 11/04
20130101; F05D 2260/22141 20130101; F01D 11/006 20130101 |
Class at
Publication: |
416/1 ;
416/97.R |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Claims
1. A turbine airfoil comprising: a platform; a blade extending from
the platform; and a slot formed in a slashface of the platform, the
slot being configured to receive a pressurized fluid via passages
and configured to direct the pressurized fluid to a selected region
of the turbine airfoil improve airfoil life.
2. The turbine airfoil of claim 1, wherein the slot is configured
to direct the pressurized fluid to lower mixing loss regions to
improve aerodynamic performance.
3. The turbine airfoil of claim 1, wherein the blade is configured
to extend into a hot gas path and the slot is configured to form a
barrier with the pressurized fluid to restrict flow of hot gas
across the slashface to a shank cavity.
4. The turbine airfoil of claim 1, wherein the slot is configured
to be joined to an adjacent slashface of an adjacent airfoil.
5. The turbine airfoil of claim 4, wherein the adjacent slashface
comprises an adjacent slot to receive the pressurized fluid from
the passages in the slashface.
6. The turbine airfoil of claim 4, wherein the passages in the
slashface are configured to provide the pressurized fluid to an
adjacent airfoil via the adjacent slashface.
7. The turbine airfoil of claim 6, wherein the adjacent slashface
comprises an adjacent slot with a passage to receive the
pressurized fluid.
8. The turbine airfoil of claim 1, wherein the slot comprises one
open end to allow the pressurized fluid to flow out from the
turbine airfoil.
9. The turbine airfoil of claim 1, wherein the slot comprises a
cross sectional geometry of one selected from the group consisting
of: a semicircle, a trapezoid and a rectangle.
10. The turbine airfoil of claim 1, comprising features in the
slashface to enable flow of the pressurized fluid on an upper
surface of the platform.
11. A turbine component assembly comprising: a first component with
passages in a body of the first component for flow of pressurized
fluid; a first slot formed in a first slashface of the body of the
first component, the first slot being configured to receive the
pressurized fluid via the passages; and a second component with a
second slashface on a body of the second component, wherein the
first slot is configured to form a barrier with the pressurized
fluid to restrict fluid communication across the first and second
slashfaces.
12. The assembly of claim 11, wherein the first slot, when joined
to the second slashface, is configured to form a barrier with the
pressurized fluid to restrict flow of hot gas across the first and
second slashfaces to improve component life by controlling a
temperature of at least one of the first and second components.
13. The assembly of claim 11, wherein the passage in the first
slashface is configured to provide the pressurized fluid to the
second component.
14. The assembly of claim 11, wherein the first component and
second component each comprise an airfoil.
15. The assembly of claim 14, wherein the second slashface
comprises a second slot.
16. The assembly of claim 14, comprising a passage in the second
slot to receive the pressurized fluid into passages within the
second component.
17. A method for controlling a temperature of a turbine airfoil,
the method comprising: flowing a pressurized fluid into a passage
formed in a platform of the turbine airfoil; and flowing the
pressurized fluid from the passage into a slot formed in a
slashface of the platform, the slot being configured to direct the
pressurized fluid to a selected region of the turbine airfoil to
improve airfoil life.
18. The method of claim 17, wherein flowing the pressurized fluid
from the passage into the slot comprises forming a barrier with the
pressurized fluid to restrict flow of hot gas across the
slashface.
19. The method of claim 17, comprising joining the slot in the
slashface to an adjacent slashface of an adjacent airfoil.
20. The method of claim 19, wherein the adjacent slashface
comprises an adjacent slot to receive the pressurized fluid from
the passages in the slashface.
Description
BACKGROUND OF THE INVENTION
[0001] The subject matter disclosed herein relates to turbines.
More particularly, the subject matter relates to an airfoil to be
positioned in a turbine.
[0002] In a gas turbine engine, a combustor converts chemical
energy of a fuel or an air-fuel mixture into thermal energy. The
thermal energy is conveyed by a fluid, often air from a compressor,
to a turbine where the thermal energy is converted to mechanical
energy. Several factors influence the efficiency of the conversion
of thermal energy to mechanical energy. The factors may include
blade passing frequencies, fuel supply fluctuations, fuel type and
reactivity, combustor head-on volume, fuel nozzle design, air-fuel
profiles, flame shape, air-fuel mixing, flame holding, combustion
temperature, turbine component design, hot-gas-path temperature
dilution, and exhaust temperature. For example, high combustion
temperatures in selected locations, such as the combustor and
turbine nozzle areas, may enable improved combustion efficiency and
power production. In some cases, high temperatures in certain
combustor and turbine regions may shorten the life and increase
wear and tear of certain components. Accordingly, it is desirable
to control temperatures in the turbine to reduce wear and increase
the life of turbine components.
BRIEF DESCRIPTION OF THE INVENTION
[0003] According to one aspect of the invention, a turbine airfoil
includes a platform and a blade extending from the platform. The
airfoil also includes a slot formed in a slashface of the platform,
the slot being configured to receive a pressurized fluid via
passages and configured to direct the pressurized fluid to a
selected region of the turbine airfoil to improve airfoil life.
[0004] According to another aspect of the invention, a method for
cooling a turbine airfoil is provided, wherein the method includes
flowing a pressurized fluid into a passage formed in a platform of
the turbine airfoil. The method also includes flowing the
pressurized fluid from the passage into a slot formed in a
slashface of the platform, the slot being configured to direct the
pressurized fluid to a selected region of the turbine airfoil to
improve airfoil life.
[0005] 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
[0006] 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:
[0007] FIG. 1 is a schematic drawing of an embodiment of a gas
turbine engine, including a combustor, fuel nozzle, compressor and
turbine;
[0008] FIG. 2 is a side view of an embodiment of an airfoil;
[0009] FIG. 3 is an end view of an embodiment of an assembly of
airfoils;
[0010] FIG. 4 is a perspective view of another embodiment of an
airfoil;
[0011] FIG. 5 is a detailed end view of an embodiment of an
airfoil; and
[0012] FIG. 6 is a detailed end view of yet another embodiment of
an airfoil.
[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] FIG. 1 is a schematic diagram of an embodiment of a gas
turbine system 100. The system 100 includes a compressor 102, a
combustor 104, a turbine 106, a shaft 108 and a fuel nozzle 110. In
an embodiment, the system 100 may include a plurality of
compressors 102, combustors 104, turbines 106, shafts 108 and fuel
nozzles 110. As depicted, the compressor 102 and turbine 106 are
coupled by the shaft 108. The shaft 108 may be a single shaft or a
plurality of shaft segments coupled together to form shaft 108.
[0015] In an aspect, the combustor 104 uses liquid and/or gas fuel,
such as natural gas or a hydrogen rich synthetic gas, to run the
turbine engine. For example, fuel nozzles 110 are in fluid
communication with a fuel supply and pressurized air from the
compressor 102. The fuel nozzles 110 create an air-fuel mix, and
discharge the air-fuel mix into the combustor 104, thereby causing
a combustion that creates a hot pressurized exhaust gas. The
combustor 104 directs the hot pressurized exhaust gas through a
transition piece into a turbine nozzle (or "stage one nozzle"),
causing turbine 106 rotation as the gas exits the nozzle or vane
and gets directed to the turbine bucket or blade. The rotation of
turbine 106 causes the shaft 108 to rotate, thereby compressing the
air as it flows into the compressor 102. In an embodiment, airfoils
(also nozzles or buckets) are located in various portions of the
turbine, such as in the compressor 102 or the turbine 106, where
hot gas flow across the airfoils causes wear and thermal fatigue of
turbine parts, due to non-uniform temperatures. Controlling the
temperature of parts of the turbine airfoil can reduce wear and
enable higher combustion temperatures in the combustor, thereby
improving performance. Controlling the temperature of regions of
and proximate to parts, such as airfoils, to improve component life
is discussed in detail below with reference to FIGS. 2-6. Although
the following discussion primarily focuses on gas turbines, the
concepts discussed are not limited to gas turbines.
[0016] FIG. 2 is a side view of a portion of an exemplary airfoil
200. The airfoil 200 includes a platform 202 and a blade 204
extending from the platform 202. A lower portion 206 extends below
the platform 202 and may be used to secure the airfoil to a part of
a rotor or stator, such as a turbine wheel. A slot 208 is formed in
a slashface 210 of the platform 202. The slashface 210 is a surface
of the platform configured to be placed adjacent to a similar
surface, or slashface, of an adjacent airfoil. A plurality of
passages 212 are located in the slot and are configured to
communicate a fluid, such as a pressurized cooling fluid or
pressurized temperature controlling fluid, into the slot 208.
Embodiments of the slashface 210 may include a single passage 212
to communicate the fluid. In an embodiment, the slashface 210 is
joined to an adjacent slashface and the pressurized fluid flows
into the slot 208 to form a fluid barrier configured to restrict
fluid flow across the slashfaces. In addition, the flow of
pressurized fluid along the slot 208 provides a distributing
cooling of the platform slashface 202, thereby reducing wear and
thermal fatigue while also improving and extending airfoil
life.
[0017] As depicted, a hot gas path 214 flows from a leading edge
216 to a trailing edge 218 of the blade 204. The pressurized fluid
barrier formed within the slot 208 restricts flow of the hot gas
across the slashface 210 to a cavity 220 (also called a "shank
cavity") in the lower portion 206. A recess 222 to receive a pin is
located below the platform 202. In embodiments, the pressurized
fluid is also configured to cool the recess 222 and pin region. By
restricting the hot gas flow across the slashface 210, the cooling
fluid within the slot 208 reduces wear and tear on the lower
portion 206. In an embodiment, the pressurized fluid is pressurized
air used to cool selected portions of the airfoil 200, wherein
passages are used to direct the cooling fluid to the selected
portions. Further, the passages may include passages 212, wherein
the pressurized fluid is distributed by the slot 208 to cool the
platform 202. In the embodiment, the slot 208 comprises a
substantially semicircular cross section geometry. As depicted, the
pressurized fluid is configured to flow in the direction of the hot
gas path 214 flow, wherein the fluid exits the open trailing edge
side of the slot 208. In other embodiments, both ends of the slot
208 may be closed. The slot 208 with closed ends may be configured
to direct the pressurized fluid to other regions of the airfoil
200. In embodiments, the slot 208 in the slashface 210 may also
provide stress relief for high stress regions of the airfoil 200,
such as the trailing edge 218 and platform 202, wherein the slot
208 weakens the slashface to divert a load from the high stress
region. As depicted, the cross sectional geometry of the slot 208
is a portion of a circle, ellipse or oval. In other embodiments,
the cross sectional geometry will include any suitable shape, such
as triangles, rectangles or trapezoids. Further, the slot 208 may
have a substantially uniform cross-section across the slashface
210. Other embodiments may have a variable cross-section for the
slot 208, such as a slot 208 that varies in cross section shape or
size along its length. For example, the slot 208 may have a
decreasing cross-section size in one direction to force flow out of
the slot 208, or with increasing size to reduce flow velocity at
the slot exit. In another example, the slot 208 could transition
from a shape optimized for heat transfer at one part of the slash
face 210 to one that is optimized for stress relief at another part
of the slash face 210.
[0018] In aspects, turbine parts, including airfoils, are formed of
stainless steel or an alloy, where the parts may experience thermal
fatigue if not properly cooled during engine operation. It should
be noted that the apparatus and method for controlling temperature
in turbine parts may apply to cooling of turbine buckets, as shown
in FIGS. 2-6, as well as nozzles, compressor vanes or any other
airfoil or hot gas path component within a turbine engine.
[0019] FIG. 3 is an end view of an exemplary assembly of an airfoil
300 and airfoil 200. The airfoil 300 is substantially similar to
airfoil 200 and includes a platform 302, a blade 304 and a lower
portion 306. The platform 302 is part of the airfoil body and
includes a slot 308 formed in a slashface 310. The slashfaces 210
and 310 are joined as the airfoils 200, 300 are assembled in a
turbine, such as on a rotor or stator. The slots 208 and 308 form a
cavity 312 that receives the pressurized fluid flow. The cavity 312
enables flow of the pressurized fluid to control the temperature of
the platforms 202 and 302. Further, the cooling fluid barrier is
formed in the cavity 312 to restrict a hot gas flow 314 across the
slashfaces 210 and 310. In the embodiment, the airfoils 200 and 300
include additional slots 316 and 318 formed in slashfaces 320 and
322, respectively. The slashfaces 320 and 322 may be joined to
slashfaces of adjacent airfoils. In an exemplary embodiment, a
passage 324 (also referred to as "channel") is located in the
airfoil 200 body and provides the pressurized fluid to the slot 208
and supplies cooling fluid flow into the slot 308 and a passage
326. Thus, the body of airfoil 200 may receive the pressurized
fluid from a source and supply the pressurized fluid to the airfoil
300 via passages 324 and 326, thereby cooling selected regions of
the airfoil 300.
[0020] FIG. 4 is a perspective view of a portion of an exemplary
airfoil 400 that includes a platform 402, a blade 404 and a lower
portion 406. The platform 402 includes a slot 408 formed in a
slashface 410 for receiving pressurized fluid from passages 412.
The platform 402 also includes features, such as notches 414, to
flow the pressurized fluid along a surface 416 of the platform 402.
Accordingly, the pressurized fluid flows 418 toward an open end of
the slot 408 and through notches 414. The pressurized fluid in the
slot 408 provides distributed cooling of the platform 402 and forms
a barrier to restrict fluid flow across the slashface 410. By
flowing the pressurized fluid through the notches 414 and to
selected regions, such as the surface 416, the slot 408 and notches
414 reduce thermal fatigue and wear. The slot 408 may include any
suitable cooling features, such as the exemplary notches 414, which
utilize structures, geometries and/or passages to direct fluid flow
onto and/or through selected portions of the airfoil, such as the
platform 402. Accordingly, by directing fluid onto the surface 416
via the notches 414, the temperature of the surface 416 region is
controlled to reduce wear and thermal fatigue. In embodiments,
cooling features may include passages and/or notches configured to
cool regions such as the blade 204, 304, 404 and/or lower portion
206, 306, 406.
[0021] FIGS. 5 and 6 are detailed end views of exemplary platforms
500 and 600 utilizing different cross sectional geometries for
slots 502 and 602, respectively. Exemplary geometries include
semi-circles, ovals, trapezoids and rectangles. The slot 502
comprises a rectangular cross sectional geometry in a slashface
504, wherein the geometry is configured to provide flow of
pressurized fluid to selected regions of the platform 500.
Similarly, the slot 602 comprises a trapezoidal cross sectional
geometry in a slashface 604. Thus, the cross sectional geometries
of the slots 208, 308, 408, 502, 602 are configured to provide
cooling to selected portions of the airfoils and/or form fluid
barriers of selected volumes to restrict fluid flow. The slots may
be formed by any suitable method, such as casting and/or machining
the platform. Further, the pressurized fluid may be provided from
an external and dedicated source, such as a coolant tank, or may be
cool air provided internally by other portions of the turbine. The
slot and suitable cross sectional geometry may be utilized for
cooling any turbine hot gas path component, wherein the slot
provides cooling and or restricts fluid flow for the component. In
an embodiment, the slot is configured to direct the pressurized
fluid to lower mixing loss regions of the airfoil to improve
aerodynamic performance. For example, the cooling fluid may be
directed to an area of the airfoil that, when it encounters other
fluid flow, such as hot gas, does not produce substantial amounts
of turbulence. In embodiments, the cooling fluid is directed to
regions of the airfoil to enable energy from the cooling fluid.
Such regions may include regions proximate the throat of the
airfoil.
[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.
* * * * *