WO2009085620A1 - Segment de buse de turbine et ensemble - Google Patents

Segment de buse de turbine et ensemble Download PDF

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Publication number
WO2009085620A1
WO2009085620A1 PCT/US2008/086310 US2008086310W WO2009085620A1 WO 2009085620 A1 WO2009085620 A1 WO 2009085620A1 US 2008086310 W US2008086310 W US 2008086310W WO 2009085620 A1 WO2009085620 A1 WO 2009085620A1
Authority
WO
WIPO (PCT)
Prior art keywords
band
tabs
turbine nozzle
leaf seal
nozzle segment
Prior art date
Application number
PCT/US2008/086310
Other languages
English (en)
Inventor
Clive Andrew Morgan
Todd Stephen Heffron
Sanjeewa Thusitha Fonseka
Original Assignee
General Electric Company
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 Company filed Critical General Electric Company
Priority to DE112008003452T priority Critical patent/DE112008003452T5/de
Priority to GB1011336A priority patent/GB2467897A/en
Priority to CA2709933A priority patent/CA2709933A1/fr
Priority to JP2010540754A priority patent/JP2011508151A/ja
Publication of WO2009085620A1 publication Critical patent/WO2009085620A1/fr

Links

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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/246—Fastening of diaphragms or stator-rings
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/55—Seals
    • F05D2240/57—Leaf seals
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • Gas turbine engines typically include a compressor, a combustor, and at least one turbine.
  • the compressor may compress air, which may be mixed with fuel and channeled to the combustor. The mixture may then be ignited for generating hot combustion gases, and the combustion gases may be channeled to the turbine.
  • the turbine may extract energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator.
  • the turbine may include a stator assembly and a rotor assembly.
  • the stator assembly may include a stationary nozzle assembly having a plurality of circumferentially spaced apart airfoils extending radially between inner and outer bands, which define a flow path for channeling combustion gases therethrough.
  • the airfoils and bands are formed into a plurality of segments, which may include one (typically called a singlet) or two spaced apart airfoils radially extending between an inner and an outer band. The segments are joined together to form the nozzle assembly.
  • the rotor assembly may be downstream of the stator assembly and may include a plurality of blades extending radially outward from a disk.
  • Each rotor blade may include an airfoil, which may extend between a platform and a tip.
  • Each rotor blade may also include a root that may extend below the platform and be received in a corresponding slot in the disk.
  • the disk may be a blisk or bladed disk, which may alleviate the need for a root and the airfoil may extend directly from the disk.
  • the rotor assembly may be bounded radially at the tip by a stationary annular shroud.
  • the shrouds and platforms define a flow path for channeling the combustion gases therethrough.
  • the nozzles and shrouds are separately manufactured and assembled into the engine. Accordingly, gaps are necessarily provided therebetween for both assembly purposes as well as for accommodating differential thermal expansion and contraction during operation of the engine.
  • the gaps between the stationary components are suitably sealed for preventing leakage therethrough.
  • a portion of air is bled from the compressor and channeled through the nozzles for cooling thereof.
  • the use of bleed air reduces the overall efficiency of the engine and, therefore, is minimized whenever possible.
  • the bleed air is at a relatively high pressure, which is greater than the pressure of the combustion gases flowing through the turbine nozzle. As such, the bleed air would leak into the flow path if suitable seals were not provided between the stationary components.
  • a typical seal used to seal these gaps is a leaf seal.
  • a typical leaf seal is arcuate and disposed end to end around the circumference of the stator components.
  • the radially outer band of the nozzle includes axially spaced apart forward and aft rails.
  • the rails extend radially outwardly and abut a complementary surface of an adjoining structural component, such as, but not limited to, a shroud, a shroud hanger, and/or a combustor liner, for providing a primary friction seal therewith.
  • the leaf seal provides a seal at this junction and bridges a portion of the rail and the adjoining structural component.
  • Leaf seals are typically relatively thin, compliant sections, which are adapted to slide along a pin fixed to one of the adjoining structural components.
  • leaf seals While leaf seals have found widespread use in turbine engines, their effectiveness in creating a fluid tight seal is dependent on the presence of a sufficient pressure differential between one side of the seal and the other. During certain operating stages of a turbine engine, the difference in fluid pressure on opposite sides of the leaf seals is relatively low. Under these conditions, it is possible for the leaf seals to unseat from their engagement with the abutting structural components of the turbo machine and allow leakage therebetween. A relatively small pressure differential across the leaf seals also permits movement or vibration of the leaf seals with respect to the structural components that they contact. This vibration of the leaf seals, which is caused by operation of the turbine engine and other sources, creates undesirable wear both of the leaf seals and the surfaces of the structural components against which the leaf seals rest. Such wear not only results in leakage of gases between the leaf seals and structural components of the turbine engine, but can cause premature failure thereof.
  • a biasing structure such as a spring
  • a band may have two circumferentially spaced apart, radially extending tabs spaced axially from a rail. A recess may be formed between the tabs and the rail where the leaf seal and spring are disposed.
  • the tabs, leaf seals and springs may include holes for receiving a pin for mounting to the band. At least one of the tabs is typically spaced apart from the circumferential edges of the band. The tab, leaf seal and spring are arranged so that the spring forces the leaf seal against an adjoining structural component so as to maintain the leaf seal in a closed, sealed position at all times.
  • low emissions combustors are susceptible to flame instability, which may lead to acoustic resonance and high dynamic pressure variation.
  • the high frequency pressure fluctuations can damage the leaf seals, particularly the leaf seals between the aft edge of the combustor liner and the leading edge of the nozzle bands, by repeatedly loading and unloading the seals against the adjoining structural component.
  • the seals are particularly susceptible to damage where they are unsupported by the springs and/or tabs. The seals may not be fully supported at their circumferential edges and/or between the tabs on the bands.
  • a turbine nozzle assembly includes a plurality of turbine nozzle segments assembled together to form an annular ring, each of the segments having an outer band having three or more circumferentially spaced apart tabs and a rail axially spaced from the tabs defining a recess therebetween.
  • the segments also have a leaf seal disposed in the recess, a pin extending through each of the tabs and the leaf seal and a biasing structure associated with each of the pins, biasing the leaf seal in abutting contact with an adjoining component.
  • the segments further include an inner band and a plurality of airfoils extending between the outer and inner bands.
  • Figure 1 is a cross-sectional schematic view of an exemplary gas turbine engine.
  • Figure 2 is a cross-sectional schematic view of an exemplary turbine nozzle assembly.
  • Figure 3 is a perspective view of an exemplary turbine nozzle segment.
  • Figure 4 is a close-up cross-sectional view of an exemplary turbine nozzle leaf seal assembly.
  • Figure 5 is a top view of an exemplary turbine nozzle segment.
  • FIG. 1 illustrates a cross-sectional schematic view of an exemplary gas turbine engine 100.
  • the gas turbine engine 100 may include a low- pressure compressor 102, a high-pressure compressor 104, a combustor 106, a high- pressure turbine 108, and a low-pressure turbine 110.
  • the low-pressure compressor may be coupled to the low-pressure turbine through a shaft 112.
  • the high-pressure compressor 104 may be coupled to the high-pressure turbine 108 through a shaft 114.
  • air flows through the low-pressure compressor 102 and high-pressure compressor 104.
  • the highly compressed air is delivered to the combustor 106, where it is mixed with a fuel and ignited to generate combustion gases.
  • the combustion gases are channeled from the combustor 106 to drive the turbines 108 and 110.
  • the turbine 110 drives the low-pressure compressor 102 by way of shaft 112.
  • the turbine 108 drives the high-pressure compressor 104 by way of shaft 114.
  • the high-pressure turbine 108 may include a turbine nozzle assembly 116.
  • the turbine nozzle assembly 116 may be downstream of the combustor 106 or a row of turbine blades.
  • the turbine nozzle assembly 116 includes an annular array of turbine nozzle segments 118.
  • a plurality of arcuate turbine nozzle segments 118 may be joined together to form an annular turbine nozzle assembly 116.
  • the nozzle segments 118 may include one or more airfoils 120 extending between an inner band 122 and an outer band 124.
  • the nozzle segments 118 may be formed in a singlet or doublet configuration.
  • the singlet configuration may include a single airfoil 120 extending between an inner band 122 and an outer band 124.
  • the inner band 122 may include a forward rail 126 and an aft rail 128.
  • the inner band 122 may also have a plurality of circumferentially spaced apart tabs 130.
  • the tabs 130 may be axially spaced from the forward rail 126 defining a recess 132 between the tabs 130 and the forward rail 126.
  • a leaf seal 134 may be disposed within the recess 132 and positioned to abut an adjoining component.
  • the adjoining component may be a combustor liner, such as combustor liner 136.
  • the adjoining component may be a turbine shroud.
  • the leaf seal 134 may be retained in the recess 132 with a pin 138.
  • the pin 138 may be positioned through a hole 140 in the tab 130 and a corresponding hole 142 in the leaf seal 134.
  • a biasing structure 144 may be retained by the pin 138 and bias the leaf seal 134 into abutting contact with the adjoining component. As shown in Figure 3, the tab 130, pin 138 and biasing structure 144, may be adjacent a circumferential edge 146 and/or a circumferential edge 147 of the nozzle segment 118.
  • the outer band 124 may include a forward rail 148 and an aft rail 150.
  • the outer band 124 may also have a plurality of circumferentially spaced apart tabs 152.
  • the tabs 152 may be axially spaced from the forward rail 148 defining a recess 154 between the tabs 152 and the forward rail 148.
  • a leaf seal 156 may be disposed within the recess 154 and positioned to abut an adjoining component.
  • the adjoining component may be a combustor liner, such as combustor liner 158.
  • the adjoining component may be a turbine shroud.
  • the leaf seal 156 may be retained in the recess 154 with a pin 160.
  • the pin 160 may be positioned through a hole 162 in the tab 152 and a corresponding hole 164 in the leaf seal 156.
  • a biasing structure 166 may be retained by the pin 160 and bias the leaf seal 156 into abutting contact with the adjoining component.
  • the tab 152, pin 160 and biasing structure 166 may be adjacent a circumferential edge 168 and/or a circumferential edge 170 of the nozzle segment 118.
  • Figures 3, 5 and 6 illustrate a nozzle segment 118 having a doublet configuration 174, where two nozzle segments 118 having a singlet configuration 172 are joined together or the nozzle segment 118 having a doublet configuration 174 is cast as one piece.
  • the airfoils 120, inner band 122 and/or outer band 124 may be formed as an integrally cast piece or may be formed separately and joined together by brazing.
  • an airfoil 120 may be integrally cast with an outer band 124 and an inner band 122 may be brazed to the airfoil.
  • the tabs 130 and 152 may be cast integrally with the inner and outer bands, respectively.
  • the outer band 124 may have a plurality of circumferentially spaced apart tabs 152, at least one of which is adjacent to a circumferential edge 168, 170 of the outer band.
  • the inner band 122 may have three or more tabs 130, one adjacent to a circumferential edge 146 of the inner band 122, one adjacent to another circumferential edge 147 of the inner band 122, and one or more therebetween.
  • the outer band 124 may have three or more tabs 152, one adjacent to a circumferential edge 168 of the outer band 124, one adjacent to another circumferential edge 170 of the outer band 124, and one or more therebetween.
  • the inner band 122 may have three or more tabs 130, one adjacent to a circumferential edge 146 of the inner band 122, one adjacent to another circumferential edge 147 of the inner band 122, and one or more therebetween.
  • the outer band 124 may also have three or more tabs 152, one adjacent to a circumferential edge 168 of the outer band 124, one adjacent to another circumferential edge 170 of the outer band 124, and one or more therebetween.
  • the leaf seals are biased into abutting contact with adjoining components to provide sealing between the turbine nozzle segment and the adjoining components.
  • the exemplary embodiments described provide additional support to the leaf seals in areas susceptible to damage, such as, but not limited to, areas adjacent to the circumferential edges of the inner and/or outer bands and the central areas therebetween.
  • the exemplary embodiments may also increase the mechanical sealing load and reduce the unsupported length of the leaf seals.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne un segment de buse de turbine comprenant une bande (122, 124) composée d'une pluralité de languettes circonférentiellement espacées (130) et une surface portante unique (120) s'étendant depuis la bande.
PCT/US2008/086310 2007-12-29 2008-12-11 Segment de buse de turbine et ensemble WO2009085620A1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
DE112008003452T DE112008003452T5 (de) 2007-12-29 2008-12-11 Turbinenleitapparatsegment und -anordnung
GB1011336A GB2467897A (en) 2007-12-29 2008-12-11 Turbine nozzle segment and assembly
CA2709933A CA2709933A1 (fr) 2007-12-29 2008-12-11 Segment de buse de turbine et ensemble
JP2010540754A JP2011508151A (ja) 2007-12-29 2008-12-11 タービンノズルセグメントおよびタービンノズルアセンブリ

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US11/967,167 US20090169369A1 (en) 2007-12-29 2007-12-29 Turbine nozzle segment and assembly
US11/967,167 2007-12-29

Publications (1)

Publication Number Publication Date
WO2009085620A1 true WO2009085620A1 (fr) 2009-07-09

Family

ID=40377633

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2008/086310 WO2009085620A1 (fr) 2007-12-29 2008-12-11 Segment de buse de turbine et ensemble

Country Status (6)

Country Link
US (1) US20090169369A1 (fr)
JP (1) JP2011508151A (fr)
CA (1) CA2709933A1 (fr)
DE (1) DE112008003452T5 (fr)
GB (1) GB2467897A (fr)
WO (1) WO2009085620A1 (fr)

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EP2299057B1 (fr) * 2009-09-04 2012-11-21 Alstom Technology Ltd Turbine à gaz
EP2336496B1 (fr) * 2009-12-14 2016-06-15 Siemens Aktiengesellschaft Moteur à turbine à gaz ayant un assemblage d'étanchéité d'aubes directrices
DE102010005153A1 (de) * 2010-01-21 2011-07-28 MTU Aero Engines GmbH, 80995 Gehäusesystem für eine Axialströmungsmaschine
US8702374B2 (en) 2011-01-28 2014-04-22 Siemens Aktiengesellschaft Gas turbine engine
CN102644484B (zh) * 2011-02-16 2016-03-23 西门子公司 燃气涡轮发动机
US8834109B2 (en) * 2011-08-03 2014-09-16 United Technologies Corporation Vane assembly for a gas turbine engine
FR2991387B1 (fr) * 2012-06-01 2016-03-04 Snecma Turbomachine, telle qu'un turboreacteur ou un turbopropulseur d'avion
US9650905B2 (en) 2012-08-28 2017-05-16 United Technologies Corporation Singlet vane cluster assembly
EP2762679A1 (fr) * 2013-02-01 2014-08-06 Siemens Aktiengesellschaft Aube de rotor de turbine à gaz et turbine à gaz
WO2014133938A1 (fr) * 2013-02-26 2014-09-04 United Technologies Corporation Renforcement de plateforme d'aube de stator de moteur à turbine à gaz
WO2015023576A1 (fr) * 2013-08-15 2015-02-19 United Technologies Corporation Panneau de protection et cadre à cet effet
WO2015105654A1 (fr) 2014-01-08 2015-07-16 United Technologies Corporation Joint de serrage pour cadre de turbine intermédiaire de turboréacteur
US9844826B2 (en) * 2014-07-25 2017-12-19 Honeywell International Inc. Methods for manufacturing a turbine nozzle with single crystal alloy nozzle segments
US20160245104A1 (en) * 2015-02-19 2016-08-25 United Technologies Corporation Gas turbine engine and turbine configurations
US10309240B2 (en) 2015-07-24 2019-06-04 General Electric Company Method and system for interfacing a ceramic matrix composite component to a metallic component
US11473437B2 (en) * 2015-09-24 2022-10-18 General Electric Company Turbine snap in spring seal
EP3181827B1 (fr) 2015-12-15 2021-03-03 MTU Aero Engines GmbH Connexion entre composants de turbomachine
WO2016179608A2 (fr) 2016-04-29 2016-11-10 Stein Seal Company Joint d'étanchéité inter-arbres avec bague d'étanchéité asymétrique
US10907491B2 (en) * 2017-11-30 2021-02-02 General Electric Company Sealing system for a rotary machine and method of assembling same
FR3128501B1 (fr) * 2021-10-25 2023-11-10 Safran Aircraft Engines Dispositif d'étanchéité à lamelle, turbomachine qui en est pourvue et aéronef correspondant
FR3139858A1 (fr) * 2022-09-15 2024-03-22 Safran Aircraft Engines Distributeur pour une turbine d’une turbomachine d’aéronef

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US6464457B1 (en) * 2001-06-21 2002-10-15 General Electric Company Turbine leaf seal mounting with headless pins
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WO2010051110A3 (fr) * 2008-10-31 2010-07-22 General Electric Company Ailette de turbine crénelée
GB2476760A (en) * 2008-10-31 2011-07-06 Gen Electric Turbine nozzle with Crenelated outer shroud flange
US8226360B2 (en) 2008-10-31 2012-07-24 General Electric Company Crenelated turbine nozzle
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Also Published As

Publication number Publication date
CA2709933A1 (fr) 2009-07-09
GB201011336D0 (en) 2010-08-18
US20090169369A1 (en) 2009-07-02
GB2467897A (en) 2010-08-18
JP2011508151A (ja) 2011-03-10
DE112008003452T5 (de) 2010-12-30

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