EP2578807A2 - Profil aérodynamique pour système de turbine - Google Patents

Profil aérodynamique pour système de turbine Download PDF

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Publication number
EP2578807A2
EP2578807A2 EP12187542.1A EP12187542A EP2578807A2 EP 2578807 A2 EP2578807 A2 EP 2578807A2 EP 12187542 A EP12187542 A EP 12187542A EP 2578807 A2 EP2578807 A2 EP 2578807A2
Authority
EP
European Patent Office
Prior art keywords
airfoil
exterior surfaces
aerodynamic contour
turbine
hot gas
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Withdrawn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
EP12187542.1A
Other languages
German (de)
English (en)
Inventor
Tyler Clark Stevenson
Andres Jose Garcia-Crespo
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 EP2578807A2 publication Critical patent/EP2578807A2/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Withdrawn legal-status Critical Current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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion

Definitions

  • the subject matter disclosed herein relates generally to turbine systems, and more particularly to airfoils for turbine systems.
  • Turbine systems are widely utilized in fields such as power generation.
  • a conventional gas turbine system includes a compressor section, a combustor section, and at least one turbine section.
  • the compressor section is configured to compress air as the air flows through the compressor section.
  • the air is then flowed from the compressor section to the combustor section, where it is mixed with fuel and combusted, generating a hot gas flow.
  • the hot gas flow is provided to the turbine section, which utilizes the hot gas flow by extracting energy from it to power the compressor, an electrical generator, and other various loads.
  • Various components such as buckets and nozzles, are typically included in the various sections of the turbine system for interacting with flows through these sections.
  • various stages of buckets and nozzles may extend into the hot gas path of the turbine section of a turbine system.
  • the portions of such components that are exposed to a hot gas path may be at risk of damage due to the high temperatures in the hot gas path.
  • such components generally require cooling.
  • improved hot gas path components would be desired in the art.
  • an improved airfoil for a hot gas path component would be advantageous.
  • the present invention resides in an airfoil for a turbine system.
  • the airfoil includes a first body having exterior surfaces defining a first portion of an aerodynamic contour of the airfoil and formed from a first material.
  • the airfoil further includes a second body having exterior surfaces defining a second portion of an aerodynamic contour of the airfoil, the second body coupled to the first body and formed from a second material having a different temperature capability than the first material.
  • the invention resides in a nozzle for a turbine section of a turbine system.
  • the nozzle includes the airfoil as described above having exterior surfaces defining an aerodynamic contour, the aerodynamic contour comprising a pressure side and a suction side extending between a leading edge and a trailing edge.
  • FIG. 1 is a schematic diagram of a gas turbine system 10. It should be understood that the turbine system 10 of the present disclosure need not be a gas turbine system 10, but rather may be any suitable turbine system 10, such as a steam turbine system or other suitable system.
  • the gas turbine system 10 may include a compressor section 12, a combustor section 14, and a turbine section 16.
  • the compressor section 12 and turbine section 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.
  • air or another suitable working fluid is flowed through and compressed in the compressor section 12.
  • the compressed working fluid is then supplied to the combustor section 14, wherein it is combined with fuel and combusted, creating hot gases of combustion. After the hot gases of combustion are flowed through the combustor section 14, they may be flowed into the turbine section 18.
  • FIG. 2 illustrates one embodiment of portions of a turbine section 18 according to the present disclosure.
  • a hot gas path 20 may be defined within the turbine section 18.
  • Various hot gas path components, such as shrouds 22, nozzles 24, and buckets 26, may be at least partially disposed in the hot gas path 20.
  • the turbine section 18 may include a plurality of buckets 26 and a plurality of nozzles 24.
  • Each of the plurality of buckets 26 and nozzles 24 may be at least partially disposed in the hot gas path 20.
  • the plurality of buckets 26 and the plurality of nozzles 24 may be disposed in one or more annular arrays, each of which may define a portion of the hot gas path 20.
  • the turbine section 16 may include a plurality of turbine stages. Each stage may include a plurality of buckets 26 disposed in an annular array and a plurality of nozzles 24 disposed in an annular array.
  • the turbine section 16 may have three stages, as shown in FIG. 2 .
  • a first stage of the turbine section 16 may include a first stage nozzle assembly 31 and a first stage buckets assembly 32.
  • the nozzles assembly 31 may include a plurality of nozzles 24 disposed and fixed circumferentially about the shaft 18.
  • the bucket assembly 32 may include a plurality of buckets 26 disposed circumferentially about the shaft 18 and coupled to the shaft 18.
  • a second stage of the turbine section 16 may include a second stage nozzle assembly 33 and a second stage buckets assembly 34.
  • the nozzles 24 included in the nozzle assembly 33 may be disposed and fixed circumferentially about the shaft 18.
  • the buckets 26 included in the bucket assembly 34 may be disposed circumferentially about the shaft 18 and coupled to the shaft 18.
  • the second stage nozzle assembly 33 is thus positioned between the first stage bucket assembly 32 and second stage bucket assembly 34 along the hot gas path 20.
  • a third stage of the turbine section 16 may include a third stage nozzle assembly 35 and a third stage bucket assembly 36.
  • the nozzles 24 included in the nozzle assembly 35 may be disposed and fixed circumferentially about the shaft 18.
  • the buckets 26 included in the bucket assembly 36 may be disposed circumferentially about the shaft 18 and coupled to the shaft 18.
  • the third stage nozzle assembly 35 is thus positioned between the second stage bucket assembly 34 and third stage bucket assembly 36 along the hot gas path 20.
  • turbine section 16 is not limited to three stages, but rather that any number of stages are within the scope and spirit of the present disclosure.
  • hot gas path components are not limited to components in turbine sections 16. Rather, hot gas path components may be components at least partially disposed in flow paths for compressor sections 12 or any other suitable sections of a system 10.
  • each hot gas path component such as each nozzle 24 and each bucket 26, may include an airfoil 40.
  • airfoils 40 of buckets 26 may extend outward from shanks 42, while airfoils 40 of nozzles 24 may extend outward from end caps 44.
  • An airfoil 40 according to the present disclosure has exterior surfaces defining a generally aerodynamic contour.
  • the exterior surfaces may define a pressure side 52 and a suction side 54 each extending between a leading edge 56 and a trailing edge 58.
  • an airfoil 40 may further include a first body 62 and a second body 64.
  • Each body 62, 64 may define a portion of the exterior surfaces of the airfoil 40.
  • the first body 62 may have exterior surfaces defining a first portion of the aerodynamic contour of the airfoil 40
  • the second body 64 has exterior surfaces defining a second portion of the aerodynamic contour of the airfoil 40.
  • the exterior surfaces of the first body 62 may define at least a portion of the pressure side 52 and suction side 54 as well as the leading edge 56, while the exterior surfaces of the second body 64 may define the trailing edge 58.
  • the exterior surfaces of the second body 64 may further define the remaining portions of the pressure side 52 and suction side 54, or these entire sides may be defined by the first body 62.
  • the first body 62 may include exterior surfaces defining any one or more of the pressure side 52, suction side 54, leading edge 56, and trailing edge 58, or any portion thereof, while the second body 64 may include exterior surfaces defining the remaining such sides and/or edges, or portions thereof.
  • the first body 62 is formed from a first material.
  • the first material may be any suitable material, such as a suitable metal or metal alloy.
  • the first material may include aluminum, nickel, iron, carbon, chromium, and/or any other suitable metal. It should further be understood that the first material is not limited to metals, but rather that the first material may be formed from any suitable material.
  • the first body 62 may define one or more cooling passages 65 therein.
  • the cooling passages 65 may be generally linear or curvilinear, and may extend through any suitable portions of the first body 62.
  • the cooling passages 65 may be generally serpentine cooling passages.
  • Such cooling passages 65 may provide convective cooling, impingement cooling, and/or any other suitable form of cooling.
  • the second body 64 of an airfoil 40 according to the present disclosure is formed from a second material.
  • the second material according to the present disclosure has a different temperature capability than the first material. Temperature capability means ability to withstand a certain temperature before failing, such that failure of the material occurs at such temperature. Thus, a second material fails at a temperature that is different from the temperature at which the first material fails. In exemplary embodiments, the second material has a temperature capability that is higher than that of the first material.
  • the second material may be any suitable metal or metal alloy, as discussed above, or a suitable non-metal material, alloy, or composite.
  • the second material is a ceramic matrix composite ("CMC") material.
  • CMC materials are designed to withstand relatively increased temperatures, such as those temperatures in a hot gas path 20 during operation of a system 10.
  • CMC materials are typically formed from ceramic fibers embedded in a ceramic matrix.
  • the fibers and/or matrix may be formed from carbon, silicon carbide, alumina, mullite, or any other suitable materials.
  • the use of CMC materials or other suitable second materials to form the second body 64 may, if the second materials have higher temperature capabilities than the first materials, advantageously allow the second body 64 to withstand the temperatures in the hot gas path 20 during operation.
  • the second body 64 includes the trailing edge 58 of the airfoil 40, which may otherwise not be adequately cooled during operation due to the lack of internal space in the airfoil 40 in which to manufacture cooling passages or other suitable cooling apparatus.
  • a body formed from a first material or second material may include other materials that are covered with a layer of the first or second material, or may be formed solely from a first or second material.
  • the second body 64 is continuous throughout a cross-sectional profile.
  • the second body 64 has a generally solid interior with no passages or other bore holes or apertures therein. In other embodiments, however, suitable passages, bore holes, or apertures may be defined through the second body 64 or portions thereof.
  • the second body 64 may be a single, unitary component. In other embodiments as shown in FIG. 3 , however, the second body 64 may comprise a plurality of second body sections 66. Each section 66 may be formed from a second material. It should be understood that the second materials utilized for each section 66 may vary or be identical. The sections 66 may be stacked on one another in the generally radial direction, as shown, or may otherwise abut one another to form the second body 64. The use of a plurality of second body sections 66 to form the second body 64 may advantageously allow for thermal growth of the second body 64 before, during, and after operation of the system 10 due to changes in temperatures in the hot gas path 20.
  • the second body 64 may be coupled to the first body 62. When coupled together, the second body 64 and first body 62 may form the generally aerodynamic contour of the airfoil 40.
  • one of the first body 62 and the second body 64 may define a recess 72, while the other of the first body 62 and the second body 64 defines a mating protrusion 74.
  • the recess 72 may be a dovetail recess
  • the protrusion 74 is a dovetail protrusion. It should be understood, however, that the present disclosure is not limited to dovetails, and rather that any suitable female and male portions may comprise the recess 72 and protrusion 74. Engagement of the protrusion 74 in the recess 72 may mate them together to couple the first body 62 to the second body 64, as shown.
  • end caps 44 or other suitable coupling apparatus may couple the second body 64 to the first body 62.
  • an end cap 44 may include openings 76 defined therein for both the first body 62 and the second body 64.
  • the first body 62 and second body 64 may be positioned relative to each other such that they are coupled together by the end cap 44, as shown. Portions of the first body 62 and second body 64 that are disposed in the openings 76 may be secured therein.
  • welds 78 may cover the openings 76 or secure cover plates 79 to the openings 76 to secure the first body 62 and second body 64 to the end cap 44.
  • the present disclosure is not limited to recesses and protrusions or end caps for coupling the second body 64 to the first body 62 of an airfoil 40. Rather, any suitable design of the bodies or suitable coupling apparatus for coupling the bodies together is within the scope and spirit of the present disclosure.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP12187542.1A 2011-10-07 2012-10-05 Profil aérodynamique pour système de turbine Withdrawn EP2578807A2 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/268,087 US20130089431A1 (en) 2011-10-07 2011-10-07 Airfoil for turbine system

Publications (1)

Publication Number Publication Date
EP2578807A2 true EP2578807A2 (fr) 2013-04-10

Family

ID=46968092

Family Applications (1)

Application Number Title Priority Date Filing Date
EP12187542.1A Withdrawn EP2578807A2 (fr) 2011-10-07 2012-10-05 Profil aérodynamique pour système de turbine

Country Status (3)

Country Link
US (1) US20130089431A1 (fr)
EP (1) EP2578807A2 (fr)
CN (1) CN103032104A (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2015191041A1 (fr) * 2014-06-10 2015-12-17 Siemens Energy, Inc. Garniture de bord de fuite pour surface portante dans un moteur à turbine à gaz
FR3159819A1 (fr) * 2024-02-29 2025-09-05 Safran Aircraft Engines Aube de stator pour une turbomachine d’aeronef
FR3159821A1 (fr) * 2024-02-29 2025-09-05 Safran Aircraft Engines Aube de stator pour une turbomachine d’aeronef

Families Citing this family (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11268526B2 (en) 2013-07-09 2022-03-08 Raytheon Technologies Corporation Plated polymer fan
WO2015006421A1 (fr) 2013-07-09 2015-01-15 United Technologies Corporation Article polymère sous encapsulation métallique
EP3019723A4 (fr) 2013-07-09 2017-05-10 United Technologies Corporation Compresseur polymère plaqué
EP3019705B1 (fr) * 2013-07-09 2019-01-30 United Technologies Corporation Revêtement à module élevé pour le raidissement local de bords de fuite d'un plan de sustentation
CA2917916A1 (fr) 2013-07-09 2015-02-05 United Technologies Corporation Coiffe polymere plaquee
WO2015075239A1 (fr) 2013-11-25 2015-05-28 Alstom Technology Ltd Ensemble pale sur la base d'une structure modulaire pour une turbomachine
US20170002661A1 (en) 2013-12-20 2017-01-05 General Electric Technology Gmbh Rotor blade or guide vane assembly
US10626740B2 (en) 2016-12-08 2020-04-21 General Electric Company Airfoil trailing edge segment

Family Cites Families (35)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2853271A (en) * 1951-06-28 1958-09-23 Eaton Mfg Co Blade structure
US3163397A (en) * 1958-01-14 1964-12-29 Daimler Benz Ag Vane construction
US3215511A (en) * 1962-03-30 1965-11-02 Union Carbide Corp Gas turbine nozzle vane and like articles
GB1040825A (en) * 1965-04-20 1966-09-01 Rolls Royce Improvements in rotor blades and/or stator blades for gas turbine engines
GB1030829A (en) * 1965-04-27 1966-05-25 Rolls Royce Aerofoil blade for use in a hot fluid stream
GB1075910A (en) * 1966-04-04 1967-07-19 Rolls Royce Improvements in or relating to blades for mounting in fluid flow ducts
US3619077A (en) * 1966-09-30 1971-11-09 Gen Electric High-temperature airfoil
US3844728A (en) * 1968-03-20 1974-10-29 United Aircraft Corp Gas contacting element leading edge and trailing edge insert
AT293148B (de) * 1969-04-28 1971-09-27 Boehler & Co Ag Geb Verfahren zur Herstellung von Turbinenschaufeln
US3650635A (en) * 1970-03-09 1972-03-21 Chromalloy American Corp Turbine vanes
US3702221A (en) * 1971-06-15 1972-11-07 Westinghouse Electric Corp Continuous shrouding-riveted construction
US4314794A (en) * 1979-10-25 1982-02-09 Westinghouse Electric Corp. Transpiration cooled blade for a gas turbine engine
US4326833A (en) * 1980-03-19 1982-04-27 General Electric Company Method and replacement member for repairing a gas turbine engine blade member
US4786234A (en) * 1982-06-21 1988-11-22 Teledyne Industries, Inc. Turbine airfoil
DE3306896A1 (de) * 1983-02-26 1984-08-30 MTU Motoren- und Turbinen-Union München GmbH, 8000 München Heissgasbeaufschlagte turbinenschaufel mit metallenem stuetzkern und umgebendem keramischen schaufelblatt
US4790721A (en) * 1988-04-25 1988-12-13 Rockwell International Corporation Blade assembly
US5197856A (en) * 1991-06-24 1993-03-30 General Electric Company Compressor stator
US5348446A (en) * 1993-04-28 1994-09-20 General Electric Company Bimetallic turbine airfoil
US5358379A (en) * 1993-10-27 1994-10-25 Westinghouse Electric Corporation Gas turbine vane
US5782607A (en) * 1996-12-11 1998-07-21 United Technologies Corporation Replaceable ceramic blade insert
DE19751129C1 (de) * 1997-11-19 1999-06-17 Mtu Muenchen Gmbh FAN-Rotorschaufel für ein Triebwerk
US6200092B1 (en) * 1999-09-24 2001-03-13 General Electric Company Ceramic turbine nozzle
DE10110102C2 (de) * 2000-12-18 2002-12-05 Deutsch Zentr Luft & Raumfahrt Rotorschaufel
US6655921B2 (en) * 2000-12-18 2003-12-02 Deutsches Zentrum Fuer Luft- Und Raumfahrt E.V. Rotor blade
US6648597B1 (en) * 2002-05-31 2003-11-18 Siemens Westinghouse Power Corporation Ceramic matrix composite turbine vane
US7014424B2 (en) * 2003-04-08 2006-03-21 United Technologies Corporation Turbine element
US7282274B2 (en) * 2003-11-07 2007-10-16 General Electric Company Integral composite structural material
US7153096B2 (en) * 2004-12-02 2006-12-26 Siemens Power Generation, Inc. Stacked laminate CMC turbine vane
US7452182B2 (en) * 2005-04-07 2008-11-18 Siemens Energy, Inc. Multi-piece turbine vane assembly
US7316539B2 (en) * 2005-04-07 2008-01-08 Siemens Power Generation, Inc. Vane assembly with metal trailing edge segment
US7393183B2 (en) * 2005-06-17 2008-07-01 Siemens Power Generation, Inc. Trailing edge attachment for composite airfoil
DE102005061673A1 (de) * 2005-12-21 2007-07-05 Rolls-Royce Deutschland Ltd & Co Kg Vorderkantenausbildung für die Verdichterschaufeln von Gasturbinentriebwerken
US20080134685A1 (en) * 2006-12-07 2008-06-12 Ronald Scott Bunker Gas turbine guide vanes with tandem airfoils and fuel injection and method of use
US8241001B2 (en) * 2008-09-04 2012-08-14 Siemens Energy, Inc. Stationary turbine component with laminated skin
US8262345B2 (en) * 2009-02-06 2012-09-11 General Electric Company Ceramic matrix composite turbine engine

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2015191041A1 (fr) * 2014-06-10 2015-12-17 Siemens Energy, Inc. Garniture de bord de fuite pour surface portante dans un moteur à turbine à gaz
FR3159819A1 (fr) * 2024-02-29 2025-09-05 Safran Aircraft Engines Aube de stator pour une turbomachine d’aeronef
FR3159821A1 (fr) * 2024-02-29 2025-09-05 Safran Aircraft Engines Aube de stator pour une turbomachine d’aeronef

Also Published As

Publication number Publication date
US20130089431A1 (en) 2013-04-11
CN103032104A (zh) 2013-04-10

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