BACKGROUND OF THE INVENTION
The field of this disclosure relates generally to components
and, more particularly, to a component for a gas turbine engine and a method of
fabricating the same.
Many known gas turbine engines include a combustion
system for mixing fuel with compressed air and igniting the mixture to produce
combustion gases. The combustion gases are directed into a turbine system to drive a
turbine into rotation, thereby driving a fan, a compressor, and/or a generator rotatably
coupled to the turbine. In some gas turbine engines (e.g., propelling gas turbine
engines on an aircraft), the combustion gases are exhausted from the turbine system
into the ambient air, thereby providing thrust for the aircraft. In some other gas
turbine engines (e.g., gas turbine engines in a combined cycle power plant), the
combustion gases are directed from the turbine system into a heat recovery steam
generator for use in producing steam.
Some known combustion systems include a plurality of
circumferentially spaced fuel nozzles that discharge fuel for use in the combustion
process. Because these fuel nozzles may discharge fuel at different rates, there can be
circumferential areas of higher combustion gas temperatures (i.e., hot streaks)
downstream of the combustion system. This can yield a substantial temperature
increase to those engine components that encounter the hot streaks. However, since
the locations of the hot streaks can be difficult to determine and can vary from engine
to engine, at least some known engines have cooling apertures formed on many
downstream engine components that do not end up being located within a hot streak
and, therefore, do not end up experiencing a temperature increase that warrants
cooling. As a result, the downstream engine components not located in the hot streaks
have been known to be excessively cooled to temperatures that are lower than desired,
and a significant amount of undesirable cooling air has therefore been known to be
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discharged into the combustion gas flow, which decreases the overall operating
efficiency of the engine. It would be useful, therefore, to have a component that
discharges cooling air only if located within a hot streak, which would facilitate
maintaining the useful life of the engine while improving the overall operating
efficiency of the engine.
BRIEF DESCRIPTION OF THE INVENTION
In one aspect, a component for a gas turbine engine 1S
provided. The component includes a cooling aperture and a plug filling at least a
portion of the cooling aperture to prevent airflow through the cooling aperture. The
plug is configured to melt at a predetermined temperature during operation of the gas
turbine engine to permit airflow through the cooling aperture.
In another aspect, a method of fabricating a component for a
gas turbine engine is provided. The method includes forming a cooling aperture in the
component and filling at least a portion of the cooling aperture with a plug that prevents
airflow through the cooling aperture. The plug is configured to melt at a predetermined
temperature during operation of the gas turbine engine to permit airflow through the
cooling aperture.
In another aspect, a gas turbine engine is provided. The gas
turbine engine includes a combustion system and a turbine system disposed downstream
of the combustion system, wherein at least one of the combustion system and the
turbine system includes a component. The component has a cooling aperture and a plug
filling at least a portion of the cooling aperture to prevent airflow through the cooling
aperture. The plug is configured to melt at a predetermined temperature during
operation of the gas turbine engine to permit airflow through the cooling aperture.
BRIEF DESCRIPTION OF THE DRAWINGS
Figure I is a schematic illustration of an exemplary gas
turbine engine;
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Figure 2 is a schematic illustration of a combustion system of
the gas turbine engine shown in Figure 1;
Figure 3 is a schematic illustration of a portion of a turbine
nozzle of the gas turbine engine shown in Figure 1;
Figure 4 is a perspective view of a segment of the turbine
nozzle shown in Figure 3; and
Figure 5 is a schematic sectional illustration of the turbine
nozzle segment shown in Figure 4.
DETAILED DESCRIPTION OF THE INVENTION
The following detailed description sets forth a component and
a method of fabricating the same by way of example and not by way of limitation.
The description should clearly enable one of ordinary skill in the art to make and use
the component, and the description sets forth several embodiments, adaptations,
variations, alternatives, and uses of the component, including what is presently
believed to be the best mode thereof. The component is described herein as being
applied to a preferred embodiment, namely a gas turbine engine. However, it is
contemplated that the component and the method of fabricating the same have general
applications in a broad range of systems and/or a variety of other commercial,
industrial, and/or consumer applications.
Figure I is a schematic illustration of an exemplary gas
turbine engine 100 including a fan system 102, a compressor system 104, a
combustion system 106, a high pressure turbine system 108, and a low pressure
turbine system 110. Figure 2 is a schematic illustration of combustion system 106. In
the exemplary embodiment, combustion system 106 includes a plurality of spacedapart,
circumferentially arranged fuel nozzles for discharging fuel during the
combustion process, namely combustion system 106 includes a first fuel nozzle 112, a
second fuel nozzle 114, a third fuel nozzle 116, a fourth fuel nozzle 118, a fifth fuel
nozzle 120, and a sixth fuel nozzle 122. In other embodiments, gas turbine engine
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100 may have any suitable number of fuel nozzles arranged in any suitable manner.
Alternatively, gas turbine engine 100 may include any suitable number of fan
systems, compressor systems, combustion systems, and/or turbine systems configured
in any suitable manner.
Figure 3 is a schematic illustration of a portion of an annular
turbine nozzle 130 of high pressure turbine system 108. In the exemplary
embodiment, turbine nozzle 130 is a stage-one nozzle of high pressure turbine system
108. In other embodiments, turbine nozzle 130 may be in any suitable stage of high
pressure turbine system 108 or low pressure turbine system 110.
In the exemplary embodiment, turbine nozzle 130 has a
plurality of turbine nozzle segments 132 that are circumferentially arranged to form
an inner band 134 and an outer band 136, with a row of spaced-apart stator vanes that
extend from inner band 134 to outer band 136, namely a first vane 140, a second vane
142, a third vane 144, a fourth vane 146, a fifth vane 148, a sixth vane 150, a seventh
vane 152, an eighth vane 154, a ninth vane 156, a tenth vane 158, and an eleventh
vane 160. As such, a first flow path 162 is defined between first vane 140 and second
vane 142; a second flow path 164 is defined between second vane 142 and third vane
144; a third flow path 166 is defined between third vane 144 and fourth vane 146; a
fourth flow path 168 is defined between fourth vane 146 and fifth vane 148; a fifth
flow path 170 is defined between fifth vane 148 and sixth vane 150; a sixth flow path
172 is defined between sixth vane 150 and seventh vane 152; a seventh flow path 174
is defined between seventh vane 152 and eighth vane 154; an eighth flow path 176 is
defined between eighth vane 154 and ninth vane 156; a ninth flow path 178 is defined
between ninth vane 156 and tenth vane 158; and a tenth flow path 180 is defined
between tenth vane 158 and eleventh vane 160. In other embodiments, turbine nozzle
130 may have any suitable number of vanes that define any suitable number of flow
paths.
Figure 4 is a perspective view of one turbine nozzle segment
132 of turbine nozzle 130. While the configuration of one exemplary turbine nozzle
segment 132 is described in more detail below, any suitable number of turbine nozzle
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segments 132 of turbine systems 108, 110 may be configured in the same manner. In
the exemplary embodiment, turbine nozzle segment 132 includes an inner band
segment 182, an outer band segment 184, and a pair of vanes (e.g., first vane 140 and
second vane 142) extending from inner band segment 182 to outer band segment 184.
In other embodiments, turbine nozzle segment 132 may have any suitable number of
vanes extending from inner band segment 182 to outer band segment 184 (e.g.,
turbine nozzle segment 132 may have a single vane, rather than a pair of vanes). In
the exemplary embodiment, each vane 140, 142 has an airfoil shape with a concave
pressure side 186 and a convex suction side 188 joined together at a leading edge 190
and a trailing edge 192. Each vane 140, 142 also includes a plurality of cooling
apertures 194 disposed on pressure side 186 and suction side 188 proximate leading
edge 190, trailing edge 192, and areas therebetween. Alternatively, vanes 140, 142
may have any suitable airfoil shape, and turbine nozzle segment 132 may have any
suitable arrangement of cooling apertures 194 (e.g., inner band segment 182 and/or
outer band segment 184 may have cooling apertures 194).
Figure 5 is a sectional view of second vane 142 through a first
cooling aperture 196 of cooling apertures 194. While the configuration of first
cooling aperture 196 of second vane 142 is described in more detail below, any
suitable component of gas turbine engine 100 (e.g., any suitable component of
combustion system 106 and/or turbine systems 108, 110, such as any suitable turbine
nozzle component, turbine shroud component, and/or turbine blade component of
turbine systems 108, 110) may have any suitable number of cooling apertures 194
configured in the same manner as first cooling aperture 196. Along the same lines,
cooling apertures 194, 196 may be of any suitable type such as, for example, film
cooling apertures, trailing edge cooling apertures, airfoil tip cooling apertures, or
platform edge cooling apertures.
In the exemplary embodiment, first cooling aperture 196
extends through concave pressure side 186 of second vane 142 such that first cooling
aperture 196 has an inlet 198 and an outlet 200. Inlet 198 is in flow communication
with an internal cooling flow passage 202 of second vane 142, and outlet 200 is in
flow communication with first flow path 162 of turbine nozzle 130. A plug 204 is
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disposed within first cooling aperture 196 to prevent airflow through first cooling
aperture 196. In the exemplary embodiment, plug 204 is located at outlet 200. In
other embodiments, plug 204 may have any suitable location along first cooling
aperture 196. In the exemplary embodiment, plug 204 is formed from a hardened
material (e.g., a braze alloy, a pure metallic element, etc.) having a predetermined
melting temperature. In other embodiments, plug 204 may be formed from any
suitable material with a predetermined melting temperature that facilitates enabling
plug 204 to function as described herein.
During operation of gas turbine engine 100, airflow through
fan system 102 is supplied to compressor system 104, and compressed air is delivered
from compressor system 104 to combustion system 106. The compressed air is mixed
with fuel from fuel nozzles 112, 114, 116, 118, 120, 122, and the combustion gases
flow from combustion system 106 into turbine nozzle 130 of high pressure turbine
system 108.
In the exemplary embodiment, because fuel nozzles 112, 114,
116, 118, 120, 122 are circumferentially spaced apart and may discharge fuel at
different rates, hotter regions ("hpt streaks") may exist in the annular combustion gas
flow into turbine nozzle 130, and these hot streaks of combustion gases would likely
be circumferentially aligned with fuel nozzles 112, 114, 116, 118, 120, 122. More
specifically, as shown in Figure 3, first flow path 162 and second flow path 164 are
circumferentially aligned with first fuel nozzle 112 and, therefore, could receive a first
hot streak of combustion gases, thereby forming a first hot streak region 206 of
turbine nozzle 130. Fifth flow path 170 and sixth flow path 172 are circumferentially
aligned with second fuel nozzle 114 and, therefore, could receive a second hot streak
of combustion gases, thereby forming a second hot streak region 208 ofturbine nozzle
130. Ninth flow path 178 and tenth flow path 180 are circumferentially aligned with
third fuel nozzle 116 and, therefore, could receive a third hot streak of combustion
gases, thereby forming a third hot streak region 210 of turbine nozzle 130. On the
other hand, third flow path 166 and fourth flow path 168 are located circumferentially
between first hot streak region 206 and second hot streak region 208 of turbine nozzle
130 and, therefore, are likely to form a first cooler region 212 of turbine nozzle 130,
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and seventh flow path 174 and eighth flow path 176 are located circumferentially
between second hot streak region 208 and third hot streak region 210 of turbine
nozzle 130 and, therefore, are likely form a second cooler region 214 of turbine
nozzle 130. In this manner, second vane 142 may be completely within first hot
streak region 206; fourth vane 146 is likely to be completely within first cooler region
212; sixth vane 150 may be completely within second hot streak region 208; eighth
vane 154 is likely to be completely within second cooler region 214; and tenth vane
158 may be completely within third hot streak region 210.
In the exemplary embodiment, the composition of plugs 204
is selected such that that the predetermined melting temperature of plugs 204 is below
the anticipated temperature of possible hot streak regions 206, 208, 210 of turbine
nozzle 130 such that plugs 204 of second vane 142, sixth vane 150, and tenth vane
158 are configured to melt during operation of gas turbine engine 100 if any of
regions 206, 208, 210 end up being hot streak regions (i.e., if any of vanes 142, 150,
158 ends up reaching the predetermined melting temperature), thereby enabling vanes
142, 150, 158 to be cooled via cooling air discharged through outlets 200 thereof.
Yet, if any of vanes 142, 150, 158 ends up not being within a hot streak region 206,
208, 210 (i.e., if any of vanes 142, 150, 158 does not end up reaching the
predetermined melting temperature), the associated plugs 204 would remain hard
enough to prevent cooling airflow through outlets 200 thereof. In this manner,
cooling air is discharged from only those vanes that reach a temperature for which
cooling is desired. In the exemplary embodiment, all plugs of turbine nozzle 130 are
made from the same material (i.e., each vane 140, 142, 144, 146, 148, 150, 152, 154,
156, 158, 160 have plugs 204 with the same predetermined melting temperature). In
other embodiments, turbine nozzle 130 may have plugs 204 made from materials
having different predetermined melting temperatures.
It should be noted that the locations of hot streak regions can
vary from engine to engine, given that each fuel nozzle of each engine may have a
different fuel discharge rate. In one example, a first engine and a second engine may
have differently located hot streak regions of the stage one nozzle. As a result, a plug
of the first engine's stage one nozzle may melt while a plug having the same
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circumferential location in the second engine's stage one nozzle may not melt. To
account for such variation in location, the above described components and methods
enable all stage one nozzle cooling apertures to be filled with plugs having a
predetermined melting temperature that is below the expected temperature of the hot
streak regions. In this manner, without the burden of anticipating the locations of the
hot streaks within each engine, cooling air can be discharged from only those engine
components that end up experiencing a temperature for which cooling is desirable,
while cooling air is not discharged from those engine components for which cooling
is not desirable.
Along the same lines, the expected temperatures of hot streak
regions can vary from engine to engine, and can even vary within a single engine. To
account for such variation in temperature, the composition of the plugs may be chosen
from a plurality of different plug compositions having different predetermined
melting temperatures to suit the engine's expected operating temperatures (e.g., the
composition of the plugs may be chosen from a first composition having a first
predetermined melting temperature, a second composition having a second
predetermined melting temperature that is higher than the first predetermined melting
temperature, and a third composition having a third predetermined melting
temperature that is higher than the first predetermined melting temperature and the
second predetermined melting temperature). In this manner, components in different
engines may be equipped with different plug compositions (e.g., a first stator vane of
a first engine's stage one nozzle may have a plug composition that is different than a
first stator vane of a second engine's stage one nozzle). Similarly, cooling apertures
in different locations in the same engine may be equipped with different plug
compositions (e.g., the cooling apertures of a first stator vane in a first engine's stage
one nozzle may have a plug composition that is different than the cooling apertures of
a second stator vane in the first engine's stage one nozzle such that the first stator
vane and the second stator vane experience cooling at different operating
temperatures). Additionally, cooling apertures in different locations on the same
engine component may be equipped with different plug compositions (e.g., a plurality
of first cooling apertures on a first stator vane may have a first plug composition
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while a plurality of second cooling apertures on the same first stator vane may have a
second plug composition such that the first cooling apertures open at a first
predetermined temperature and the second cooling apertures open at a second
predetermined temperature, thereby providing stepwise cooling of the first stator
vane).
The methods and systems described herein facilitate
providing a gas turbine engine with cooling apertures for cooling engine components.
The methods and systems described herein further facilitate configuring the cooling
apertures of the gas turbine engine such that cooling air is provided only to those
engine components for which cooling is desired. The methods and systems described
herein also facilitate accounting for variation in the locations and temperatures of hot
streaks by filling cooling apertures with plugs that are configured to melt only when a
predetermined temperature threshold is met, thereby preventing cooler areas of the
gas turbine engine from being excessively and undesirably cooled. The methods and
systems described herein therefore facilitate maintaining the useful life of the engine
by cooling components for which cooling is desired, while improving the overall
operating efficiency ofthe engine by preventing the excessive discharge of cooling air
that results from cooling engine components for which cooling is not desired.
Exemplary embodiments of a component and a method of
fabricating the same are described above in detail. The methods and systems are not
limited to the specific embodiments described herein, but rather, components of the
methods and systems may be utilized independently and separately from other
components described herein. For example, the methods and systems described
herein may have other industrial and/or consumer applications and are not limited to
practice with only gas turbine engines as described herein. Rather, the present
invention can be implemented and utilized in connection with many other industries.
While the invention has been described in terms of various
specific embodiments, those skilled in the art will recognize that the invention can be
practiced with modification within the spirit and scope ofthe claims.
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PARTS LIST
100 gas turbine engine
102 fan system
104 compressor system
106 combustion system
108 high pressure turbine system
110 low pressure turbine system
112 first fuel nozzle
114 second fuel nozzle
116 third fuel nozzle
118 fourth fuel nozzle
120 fifth fuel nozzle
122 sixth fuel nozzle
130 turbine nozzle
132 turbine nozzle segment
134 inner band
136 outer band
140 first vane
142 second vane
144 third vane
146 fourth vane
148 fifth vane
150 sixth vane
152 seventh vane
154 eighth vane
156 ninth vane
158 tenth vane
160 eleventh vane
, I
251262
162 first flow path
164 second flow path
166 third flow path
168 fourth flow path
170 fifth flow path
172 sixth flow path
174 seventh flow path
176 eighth flow path
178 ninth flow path
180 tenth flow path
182 inner band segment
184 outer band segment
186 concave pressure side
188 convex suction side
190 leading edge
192 trailing edge
194 cooling apertures
196 first cooling aperture
198 inlet
200 outlet
202 internal cooling flow passage
204 plug
206 first hot streak region
208 second hot streak region
210 third hot streak region
212 first cooler region
214 second cooler region

WE CLAIM:
1. A component for a gas turbine engine (100), said component
comprising:
a cooling aperture (194, 196); and
a plug (204) filling at least a portion of said cooling aperture to prevent airflow
through said cooling aperture, wherein said plug is configured to melt at a
predetermined temperature during operation of the gas turbine engine to permit
airflow through said cooling aperture.
2. A component in accordance with Claim I, wherein said component is
one of a turbine nozzle component, a turbine shroud component, and a turbine blade
component.
3. A component in accordance with Claim 1, wherein said cooling
aperture (194, 196) is one of a film cooling aperture, a trailing edge cooling aperture,
an airfoil tip cooling aperture, and a platform edge cooling aperture.
4. A component in accordance with Claim I, wherein said cooling
aperture (194, 196) comprises an inlet (198) and an outlet (200), said plug (204)
disposed at said outlet.
5. A component in accordance with Claim I, wherein said plug (204) is
formed from a hardened metallic material.
6. A component in accordance with Claim I, wherein said plug (204) is
formed from a hardened braze alloy material.
7. A gas turbine engine (100) comprising:
a combustion system (106); and
a turbine system (108, 110) disposed downstream of said combustion system,
wherein at least one of said combustion system and said turbine system comprises a
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component comprising a cooling aperture (194, 196) and a plug (204) filling at least a
portion of said cooling aperture to prevent airflow through said cooling aperture,
wherein said plug is configured to melt at a predetermined temperature during
operation of said gas turbine engine to permit airflow through said cooling aperture.
8. A gas turbine engine (100) in accordance with Claim 7, wherein said
component is a stator vane (140, 142, 144, 146, 148, 150, 152, 154, 156, 158, 160)
comprising a convex suction side (188), a concave pressure side (186), and an internal
cooling flow passage (202), said cooling aperture (194, 196) extending through one of
said convex suction side and said concave pressure side such that said cooling
aperture is in flow communication with said cooling flow passage.
9. A gas turbine engine (100) in accordance with Claim 7, wherein said
cooling aperture (194, 196) comprises an inlet (198) and an outlet (200), said plug
(204) disposed at said outlet.
10. A gas turbine engine (100) in accordance with Claim 7, wherein said
plug (204) is formed from a hardened braze alloy material.
J~~ L..j. 0<4<.
MANISHA SINGH NAIR
Agent for the Applicant [IN/PA-740]
LEX ORBIS
Intellectual Property Practice
709/710, Tolstoy House,
15-17, Tolstoy Marg,
New Delhi-II 000 I