EP2581557A2 - Composant de partie chaude pour système de turbine - Google Patents

Composant de partie chaude pour système de turbine Download PDF

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Publication number
EP2581557A2
EP2581557A2 EP12179575.1A EP12179575A EP2581557A2 EP 2581557 A2 EP2581557 A2 EP 2581557A2 EP 12179575 A EP12179575 A EP 12179575A EP 2581557 A2 EP2581557 A2 EP 2581557A2
Authority
EP
European Patent Office
Prior art keywords
hot gas
gas path
path component
shell
interior surface
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP12179575.1A
Other languages
German (de)
English (en)
Inventor
Benjamin Paul Lacy
Brian Gene Brzek
Rebecca Christine Malish
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
General Electric Co
Original Assignee
General Electric Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by General Electric Co filed Critical General Electric Co
Publication of EP2581557A2 publication Critical patent/EP2581557A2/fr
Withdrawn legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/182—Transpiration cooling
    • F01D5/183—Blade walls being porous

Definitions

  • the subject matter disclosed herein relates generally to turbine systems, and more specifically to hot gas path components for turbine systems.
  • Turbine systems are widely utilized in fields such as power generation.
  • a conventional gas turbine system includes a compressor, a combustor, and a turbine.
  • various components in the system are subjected to high temperature flows, which can cause the components to fail. Since higher temperature flows generally result in increased performance, efficiency, and power output of the gas turbine system, the components that are subjected to high temperature flows must be cooled to allow the gas turbine system to operate at increased temperatures, increased efficiency, and/or reduced emissions.
  • a cooling medium may be routed from the compressor and provided to various components.
  • the cooling medium may be utilized to cool various compressor and turbine components.
  • Nozzles are one example of a hot gas path component that must be cooled.
  • various parts of the nozzle such as the airfoil, are disposed in a hot gas path and exposed to relatively high temperatures, and thus require cooling.
  • One solution for cooling a nozzle is to include an impingement sleeve inside the airfoil. Cooling medium is flowed to the interior of the nozzle, and then flowed through the impingement sleeve and onto an interior surface of the airfoil. This approach facilitates impingement cooling of the airfoil.
  • impingement sleeves do provide adequate cooling of nozzles, increased cooling efficiency is desired. Such increased efficiency would allow for a reduction in the cooling medium required to cool the nozzles, and thus a reduction in emission and/or increase in firing temperature.
  • an improved hot gas path component such as an improved nozzle, for a turbine system is desired in the art.
  • a hot gas path component with improved cooling features would be advantageous.
  • the present invention resides in a hot gas path component for a turbine system.
  • the hot gas path component includes a shell having an exterior surface and an interior surface.
  • the hot gas path component further includes a porous medium having an exterior surface and an interior surface, the exterior surface positioned adjacent to the interior surface of the shell.
  • the porous medium is configured for flowing a cooling medium therethrough.
  • FIG. 1 is a schematic diagram of a gas turbine system 10.
  • the system 10 may include a compressor 12, a combustor 14, and a turbine 16.
  • the compressor 12 and turbine 16 may be coupled by a shaft 18.
  • the shaft 18 may be a single shaft or a plurality of shaft segments coupled together to form shaft 18.
  • the turbine 16 may include a plurality of turbine stages.
  • the turbine 16 may have three stages.
  • a first stage of the turbine 16 may include a plurality of circumferentially spaced nozzles and buckets.
  • the nozzles may be disposed and fixed circumferentially about the shaft 18.
  • the buckets may be disposed circumferentially about the shaft and coupled to the shaft 18.
  • a second stage of the turbine 16 may include a plurality of circumferentially spaced nozzles and buckets.
  • the nozzles may be disposed and fixed circumferentially about the shaft 18.
  • the buckets may be disposed circumferentially about the shaft 18 and coupled to the shaft 18.
  • a third stage of the turbine 16 may include a plurality of circumferentially spaced nozzles and buckets.
  • the nozzles may be disposed and fixed circumferentially about the shaft 18.
  • the buckets may be disposed circumferentially about the shaft 18 and coupled to the shaft 18.
  • the various stages of the turbine 16 may be at least partially disposed in the turbine 16 in, and may at least partially define, a hot gas path. It should be understood that the turbine 16 is not limited to three stages, but rather that any number of stages are within the scope and spirit of the present disclosure.
  • the compressor 12 may include a plurality of compressor stages (not shown). Each of the compressor 12 stages may include a plurality of circumferentially spaced nozzles and buckets.
  • the hot gas path component 30 is a nozzle.
  • a hot gas path component 30 may be a bucket, a shroud block, or any other suitable component that may be disposed in the path of hot gases flowing through a turbine system 10.
  • the nozzle 30 may include a shell 32.
  • the shell 32 may be an airfoil that extends between end caps 34.
  • the shell 32 may have a generally aerodynamic contour.
  • the shell 32 may have an exterior surface 36 and an interior surface 38.
  • the exterior surface 36 may define a pressure side 42 and suction side 44 each extending between a leading edge 46 and a trailing edge 48, or any other suitable aerodynamic contour.
  • One or more of the end caps 34 may define an opening (not shown). The opening may allow cooling medium 50 to flow to the interior 52 of the shell 32, def-med by the interior surface 38, as is generally known in the art.
  • the hot gas path component 30 may further include an impingement sleeve 60, as shown in FIGS. 3 , 4 , 6 and 7 .
  • the impingement sleeve 60 may be disposed at least partially within the interior 52 of the shell 32, and spaced from the interior surface 38.
  • the impingement sleeve may have an exterior surface 62 and interior surface 64, and may have a contour similar to that of the shell 32.
  • the impingement sleeve 60 may define one or more impingement passages 66 extending between the interior surface 64 and the exterior surface 62. Cooling medium 50 flowed into the interior 52 of the shell 32 may be flowed through these impingement passages 66.
  • the hot gas path component 30 may include any suitable sleeve therein.
  • a sleeve may include a plurality of spaced apart plates which allow cooling medium 50 to flow therebetween.
  • a hot gas path component 30 further includes one or more porous media 70.
  • a porous medium 70 according to the present disclosure has an exterior surface 72 and an interior surface 74. The exterior surface 72 is positioned adjacent the interior surface 38 of the shell 32.
  • the porous media 70 are positioned between the hot gas path component 30 and impingement sleeve 60 or other suitable sleeve, such that the exterior surface 62 of the impingement sleeve 60 is positioned adjacent the interior surfaces 64 of the porous media 70.
  • the porous media 70 may advantageously allow improved cooling of the hot gas path component 30, such as of the shell 32.
  • the porous media 70 allow for conductive heat transfer from the shell 32 due to the cooling medium 50 flowing generally through the porous media 70.
  • the porous media 70 may additionally allow for impingement cooling of the shell 32, thus further improving cooling of the hot gas path component 30.
  • a porous medium 70 may be formed from any suitable porous material or materials having a matrix 76 and one or more voids 78.
  • a porous medium 70 such as the matrix 76 thereof, may be formed from a metal or metal alloy foam, a ceramic foam, such as a ceramic matrix composite foam, or a carbon fiber foam.
  • a foam is typically formed by mixing a material, such as a metal, ceramic, or carbon fiber, with another substance and then melting the substance away, leaving a porous foam.
  • the porous medium 70 may be formed from, for example, a plurality of packed together beads of a suitable material, or any other suitable material or materials.
  • the porous medium 70 may thus be configured for flowing cooling medium 50 therethrough.
  • the cooling medium 50 may flow through the voids 78 in a porous medium 70 before contacting the interior surface 38 of the shell 32, thus in exemplary embodiments facilitating convection cooling.
  • the hot gas path component 30 may include one porous medium 70.
  • the porous medium 70 is continuous in the direction of the contour, such as the aerodynamic contour, of the shell 32, such that substantially all of a cross-sectional profile of the interior surface 38 is adjacent to the porous medium 70. In other embodiments, only a portion of a cross-sectional profile of the interior surface 38 may be adjacent to the porous medium.
  • the hot gas path component 30 may include more than one porous medium 70.
  • Each of the plurality of porous mediums 70 may be spaced apart from others of the plurality of porous mediums 70, such as in the direction of the contour, such as the aerodynamic contour, of the shell 32 as shown or in any other suitable direction, or may abut or otherwise contact others of the plurality of porous mediums 70.
  • an impingement sleeve 60 may be positioned adjacent the interior surface 74 of a porous medium 70.
  • cooling medium 50 may be flowed through the impingement passages 66 of the impingement sleeve 60 to the porous medium 70.
  • no impingement sleeve 60 may be included in the hot gas path component 30.
  • the interior surfaces 74 of the porous media 70 may be treated. Such treating may seal the interior surface 74, such that voids 78 defined in a porous medium 70 do not extend to the interior surface 74. Passages, such as impingement passages, may then be formed through such treated interior surface 74, as discussed below, to allow cooling medium 50 to flow therethrough. Treating of the interior surface 74 may include grinding, filling, brazing, welding, soldering, or any other suitable treating technique that would suitably seal the interior surface 74.
  • a porous medium 70 may be in contact with the shell 32 and/or optional impingement sleeve 60.
  • the exterior surface 72 of the porous medium 70 may contact the interior surface 38 of the shell 32.
  • the interior surface 74 of the porous medium 70 may contact the exterior surface 62 of the impingement sleeve 60.
  • the porous medium 70 may be press-fit, bonded such as through a suitable adhesive or bonding process, or otherwise connected to the shell 32 and/or impingement sleeve 60.
  • a porous medium 70 may be spaced from the shell 32 and/or the impingement sleeve 60.
  • a porous medium 70 according to the present disclosure may be in contact with both an shell 32 and an impingement sleeve 60, may be spaced from both a shell 32 and an impingement sleeve 60, or may be in contact with one of an shell 32 or an impingement sleeve 60 and spaced from the other of an shell 32 or an impingement sleeve 60.
  • one or more impingement passages 80 may be defined in a porous medium 70.
  • the impingement passages 80 may extend between the interior surface 74 and the exterior surface 72 of the porous medium 70.
  • Such impingement passages 80 may allow for cooling medium 50 to flow therethrough and impinge on the inner surface 38 of the shell 32, thus impingement cooling the shell 32.
  • portions of the cooling medium 50 may enter the impingement passages 80 and then flow from the impingement passages 80 through the voids 78 in the porous medium 70, thus otherwise facilitating cooling of the shell 32.
  • Such impingement passages 80 may have any suitable cross-sectional shape, such as circular or oval-shaped, square or rectangle shaped, triangular, or having any other suitable polygonal shape.
  • the impingement passages 80 may have generally circular cross-sectional shapes, while in others the impingement passages 80 may have generally rectangular cross-sectional shapes and be characterized as slots.
  • the impingement passages 80 may have cross-sectional areas that are larger than, identical to, or smaller than those of the impingement passages 66.
  • impingement passages 80 may have any suitable cross-sectional area, and this cross-sectional area may be constant throughout the length of the passage 80 or may vary.
  • a passage 80 may taper, or may have a constricted portion or a relatively larger portion.
  • impingement passages 80 may be linear, curvilinear, or have any other suitable path.
  • an impingement passage 80 may be curvilinear, having a generally serpentine path. In other embodiments, an impingement passage 80 may simply have a linear path.
  • An impingement passage 80 may be drilled or otherwise formed into a porous medium 70.
  • the impingement passages 66 in the impingement sleeve 60 may generally align with the impingement passages 80 of the porous medium 70.
  • the impingement passages 80 may extend through this treated surface.
  • a shell 32 according to the present disclosure may further define one or more cooling passage 82, as shown in FIGS. 7 and 8 .
  • the cooling passages 80 may extend between the interior surface 38 and the exterior surface 36 of the shell 32.
  • Such cooling passages 80 may have any suitable cross-sectional shape, cross-sectional area, and cross-sectional path, as discussed above.
  • the cooling passages 80 may be film cooling passages, and may be angled and formed such that cooling medium 50 flowed therethrough and exhausted therefrom then provides film cooling to the exterior surface 36 of the shell 32.
  • a cooling passage 82 may be aligned with a porous medium 70, as shown, or with an impingement passage 80 defined therein. Cooling medium 50 flowing through the impingement passages 80 and porous medium 70 may flow into and through the cooling passage 82.
  • a cooling passage 82 extends only through the shell 32 between the interior surface 38 and exterior surface 36.
  • a cooling passage 82 may further extend at least partially into and be at least partially defined in a porous medium 70.
  • a cooling passage 82 may extend through the exterior surface 72 of a porous medium 70, as shown.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP12179575.1A 2011-10-12 2012-08-07 Composant de partie chaude pour système de turbine Withdrawn EP2581557A2 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/271,724 US20130094971A1 (en) 2011-10-12 2011-10-12 Hot gas path component for turbine system

Publications (1)

Publication Number Publication Date
EP2581557A2 true EP2581557A2 (fr) 2013-04-17

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP12179575.1A Withdrawn EP2581557A2 (fr) 2011-10-12 2012-08-07 Composant de partie chaude pour système de turbine

Country Status (3)

Country Link
US (1) US20130094971A1 (fr)
EP (1) EP2581557A2 (fr)
CN (1) CN103046973A (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3222814A1 (fr) * 2016-03-24 2017-09-27 Siemens Aktiengesellschaft Aube, procédé de fabrication associé et turbomachine associée
WO2017196498A1 (fr) * 2016-05-12 2017-11-16 General Electric Company Paroi de composant de moteur avec un circuit de refroidissement
EP3249159A1 (fr) * 2016-05-23 2017-11-29 Siemens Aktiengesellschaft Aube de turbine et turbomachine associée
WO2019141755A1 (fr) * 2018-01-18 2019-07-25 Siemens Aktiengesellschaft Concept de refroidissement pour composant de turbine
FR3115816A1 (fr) * 2020-11-05 2022-05-06 Safran Composant pour turbomachine a refroidissement ameliore

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US9663404B2 (en) * 2012-01-03 2017-05-30 General Electric Company Method of forming a ceramic matrix composite and a ceramic matrix component
US9896943B2 (en) * 2014-05-12 2018-02-20 Honeywell International Inc. Gas path components of gas turbine engines and methods for cooling the same using porous medium cooling systems
US10598026B2 (en) * 2016-05-12 2020-03-24 General Electric Company Engine component wall with a cooling circuit
US11697994B2 (en) * 2020-02-07 2023-07-11 Raytheon Technologies Corporation CMC component with cooling protection
US11746660B2 (en) 2021-12-20 2023-09-05 Rolls-Royce Plc Gas turbine engine components with foam filler for impact resistance
US11834956B2 (en) 2021-12-20 2023-12-05 Rolls-Royce Plc Gas turbine engine components with metallic and ceramic foam for improved cooling

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Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3222814A1 (fr) * 2016-03-24 2017-09-27 Siemens Aktiengesellschaft Aube, procédé de fabrication associé et turbomachine associée
WO2017196498A1 (fr) * 2016-05-12 2017-11-16 General Electric Company Paroi de composant de moteur avec un circuit de refroidissement
EP3249159A1 (fr) * 2016-05-23 2017-11-29 Siemens Aktiengesellschaft Aube de turbine et turbomachine associée
WO2019141755A1 (fr) * 2018-01-18 2019-07-25 Siemens Aktiengesellschaft Concept de refroidissement pour composant de turbine
FR3115816A1 (fr) * 2020-11-05 2022-05-06 Safran Composant pour turbomachine a refroidissement ameliore

Also Published As

Publication number Publication date
US20130094971A1 (en) 2013-04-18
CN103046973A (zh) 2013-04-17

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