EP3075959A1 - Turbine à gaz comprenant une chambre de combustion avec une sortie de chambre de combustion et une première rangée d'aubes basculantes - Google Patents

Turbine à gaz comprenant une chambre de combustion avec une sortie de chambre de combustion et une première rangée d'aubes basculantes Download PDF

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Publication number
EP3075959A1
EP3075959A1 EP15161916.0A EP15161916A EP3075959A1 EP 3075959 A1 EP3075959 A1 EP 3075959A1 EP 15161916 A EP15161916 A EP 15161916A EP 3075959 A1 EP3075959 A1 EP 3075959A1
Authority
EP
European Patent Office
Prior art keywords
gas turbine
rocking
vane
platform
vanes
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Withdrawn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
EP15161916.0A
Other languages
German (de)
English (en)
Inventor
Hans-Christian Mathews
Fabien Fleuriot
Jost Imfeld
Urs Benz
Frank Graf
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.)
Ansaldo Energia Switzerland AG
Original Assignee
Alstom Technology AG
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 Alstom Technology AG filed Critical Alstom Technology AG
Priority to EP15161916.0A priority Critical patent/EP3075959A1/fr
Publication of EP3075959A1 publication Critical patent/EP3075959A1/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Withdrawn legal-status Critical Current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
    • 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
    • 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
    • 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
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F05D2220/321—Application in turbines in gas turbines for a special turbine stage
    • F05D2220/3212—Application in turbines in gas turbines for a special turbine stage the first stage of a turbine
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/60—Assembly methods
    • F05D2230/64—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins
    • F05D2230/642—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins using maintaining alignment while permitting differential dilatation
    • 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
    • F05D2260/00—Function
    • F05D2260/94—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF]
    • F05D2260/941—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF] particularly aimed at mechanical or thermal stress reduction

Definitions

  • the present invention relates to the technology of gas turbines. It refers to a gas turbine comprising a combustor with a combustor outlet and a first row of rocking vanes according to the preamble of claim 1.
  • Fig. 1 shows part of a gas turbine 40 according to the prior art at the transition between the combustor 35 of the turbine and the first row of rocking vanes 42 at the entrance of hot gas flow path 41.
  • Hot gas 36 exits the combustor 35 through a plurality of openings, which are positioned around the machine axis in an annular arrangement and which are known as picture frames 35a.
  • Each rocking vane 42 comprises a radially oriented airfoil 43 extending between an outer platform 44 (outer diameter OD) and an inner platform 45 (inner diameter ID) and having a leading edge 43a and a trailing edge 43b. Vane 42 is hooked with its outer platform 44 into a turbine vane carrier 47. Inner platform 45 abuts a rotor cover 48 at a contact point 39, which covers the central rotor part of the machine. Outer platform 44 and inner platform 45 border the hot gas flow path 41 and are in sealing relation to picture frame 35a by means of seals 37 and 38.
  • each rocking vane 42 is loaded with a very high force under extreme boundary conditions (high temperature and high pressure). Therefore, modifications on the vane need to be carried out in a way that the lifetime of the vane on one hand and the hot gas flow after the vane on the other hand is not negatively affected.
  • the retention flange includes a radial lug trapped axially in the retention slot, and a tangential lug at an opposite circumferential end disposed on the forward flange.
  • the aft flange includes an outer hinge, and the retention flange further includes an inner hinge for cradle mounting the retention flange in the nozzle support.
  • Document US 7,762,766 B2 is related to a support for the first row turbine vanes in a gas turbine engine, the support including a support framework.
  • the support framework includes an outer vane carrier, and an inner vane carrier connected to the outer vane carrier by a plurality of struts.
  • the outer vane carrier is mounted to an inner casing of the engine and the struts support the inner vane carrier in cantilevered relation to the outer vane carrier.
  • An aft outer flange of each first row vane is supported on the outer vane carrier and a forward inner flange of the vane is supported on the inner vane carrier.
  • the gas turbine according to the invention comprises a combustor with a combustor outlet and a first row of rocking vanes arranged at said combustor outlet in sealing relationship with said combustor outlet, whereby said rocking vanes each comprise an airfoil with a leading edge and a trailing edge extending between an inner platform and an outer platform, and whereby said rocking vanes are each hooked with their outer platform into a turbine vane carrier thereby defining a rotation point for a rotational movement of said rocking vanes.
  • the gas turbine according to the invention is characterized in that said rotation point is located at the leading edge of said rocking vanes.
  • An embodiment of the gas turbine according to the invention is characterized in that a means for limiting said rotational movement of said rocking vanes is provided on said outer platform at said trailing edge.
  • said means for limiting said rotational movement comprises a stop collar on the outer side of said outer platform.
  • a defined gap is provided between said stop collar and said turbine vane carrier.
  • said gas turbine comprises a rotor with a rotor cover, and said rocking vanes are supported at said inner platform by said rotor cover.
  • said gas turbine comprises a central rotor part and a plurality of I-beams extending in radial direction between said turbine vane carrier and said central rotor part, and said rocking vanes are supported at said inner platform by said I-beams.
  • a seal especially honeycomb seal, is provided between said inner platform and said I-beam.
  • a horizontal seal especially honeycomb seal, is provided between said outer platform and said turbine vane carrier.
  • said gas turbine comprises a rotor with a rotor cover, and a seal, especially dog-bone seal, is provided between said inner platform and said rotor cover.
  • the basic idea of the present invention is to modify the rocking vane(s) in a way so that the rocking vane movements and the axial force on the ID (inner platform) are reduced (whereas some benefits of the rocking vane will be kept).
  • rocking vane can be used for the single rocking vane (support on the ID or inner platform with the rotor cover - see Fig. 2 ) or for the rocking vane together with an I-beam (support on the ID or inner platform with the I-beam - see Fig. 3 ).
  • Main features of the modified rocking vane are the change in position of the rotation hook (rotation point) and a stop collar on the OD or outer platform either with or without an additional honeycomb seal.
  • the starting point of the present invention is the gas turbine engine in closed condition, which has a pattern of several picture frames (35a in Fig. 1 ) or outlets of a sequential liner of a sequential combustor) spread around a rotational contour.
  • a main target is to seal the interface between each picture frame and the 1 st vane (in this case rocking vane) in order to reduce the leakage to a minimum.
  • a first fixation concept is shown with gas turbine 10 in Fig. 2 .
  • a single rocking vane 12 which comprises an airfoil 13 with leading edge 13a and trailing edge 13b, extending radially through hot gas flow path 11, and an inner platform (or ID) 15 and outer platform (or OD) 14, is supported at its inner platform 15 by a rotor cover 18.
  • the modified rocking vane 12 is now hooked with its outer platform 14 into turbine vane carrier (TVC) 17 at the leading edge 13a.
  • TVC turbine vane carrier
  • it has a change in position of the rotation point 19 on the outer platform 14 (with a different rotation radius RR) in order to minimize the rocking vane movement and the axial force on the vane.
  • the axial force on inner platform (ID) 15 will be reduced due to the additional moment over the vane, which results because of the different sizes of the inner and outer vane platform 15 and 14, respectively.
  • rocking vane 12 is supported at the inner platform (ID) 15 by the rotor cover 18.
  • a stop collar 16 on the outer platform (OD) 14 with a defined gap between stop collar 16 and turbine vane carrier 17 is provided in order to limit the rocking vane movement up to a defined maximum. If the maximum is reached the force on the vane will be directed into the turbine vane carrier (TVC) 17.
  • a second fixation concept is shown with gas turbine 20 in Fig. 3 .
  • a single rocking vane 22 which comprises an airfoil 23 with leading edge 23a and trailing edge 23b, extending radially through hot gas flow path 21, and an inner platform (or ID) 25 and outer platform (or OD) 24, has a support at inner platform 25 by means of an I-beam 33 (a radial strut between rotor cover and turbine vane carrier of the turbine)
  • the modified rocking vane 22 has a change in position of the rotation point 29 at outer platform (OD) 24 in order to minimize the axial force on the rocking vane 22 (see rotation radius RR).
  • a honeycomb seal 32 between inner platform (ID) 25 and I-beam 33 is provided to minimize the leakage in general and in case of vane twist.
  • a stop collar 26 on the outer platform (OD) 24 of the vane with a defined gap is provided in order to limit the rocking vane movement up to a defined maximum.
  • I-beam 33 will be relieved.
  • the leakage between rocking vane outer platform (OD) 24 and turbine vane carrier (TVC) 27 is minimized by a honeycomb seal 30.
  • Sealing between 1 st vane and rotor cover 28 is realized with an additional seal, e.g. a dog-bone seal 31.
  • FIG. 4 A third fixation concept in a gas turbine is shown in Fig. 4 .
  • the modified rocking vane 52 is hooked with its outer platform 54 into the turbine vane carrier 57 at the leading edge 53a.
  • it has a change in position of the rotation point 59 on the outer platform 54 (with a different rotation radius RR) in order to minimize the rocking vane 52 movements and the axial force on the vane.
  • rocking vane 52 is supported on the leading edge 53a at the inner platform (ID) 55 by the rotor cover 58. Therefore, the rotational point 59 of the rocking vane 52 and the support on the inner platform 55 of the vane are in line.
  • a stop collar 56 on the outer platform (OD) 54 with a defined gap between stop collar 56 and turbine vane carrier 57 is provided in order to limit the rocking vane 52 movement up to a defined maximum. If the maximum is reached the force on the vane will be directed into the turbine vane carrier 57.
  • the sealing concept of this version is based on two seals, preferably honeycomb seals 30, 32 on the outer platform OD 54 and the inner platform ID 55 of the vane 52 at its trailing edge 53b.
  • One seal 32 is located on the rotor cover 58 and is able to capture the axial and radial movements of the vane 52 and the rotor cover 58 itself.
  • the second seal 30 is placed on the turbine vane carrier 57 and prevents leakage between the vane outer platform 54 and the hot gas flow.
  • the modified rocking vane enables to reduce the forces on the inner platform and the vane movements. This has a positive impact on the leakage at the ID (leakage reduction) and improves also other sealing concepts, e.g. honeycomb seal or I-beam in terms of leakage due to the reduced vane movement.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP15161916.0A 2015-03-31 2015-03-31 Turbine à gaz comprenant une chambre de combustion avec une sortie de chambre de combustion et une première rangée d'aubes basculantes Withdrawn EP3075959A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP15161916.0A EP3075959A1 (fr) 2015-03-31 2015-03-31 Turbine à gaz comprenant une chambre de combustion avec une sortie de chambre de combustion et une première rangée d'aubes basculantes

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP15161916.0A EP3075959A1 (fr) 2015-03-31 2015-03-31 Turbine à gaz comprenant une chambre de combustion avec une sortie de chambre de combustion et une première rangée d'aubes basculantes

Publications (1)

Publication Number Publication Date
EP3075959A1 true EP3075959A1 (fr) 2016-10-05

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EP15161916.0A Withdrawn EP3075959A1 (fr) 2015-03-31 2015-03-31 Turbine à gaz comprenant une chambre de combustion avec une sortie de chambre de combustion et une première rangée d'aubes basculantes

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3082555A1 (fr) * 2018-06-18 2019-12-20 Safran Aircraft Engines Roue d'aubes redresseuses d'un moteur d'aeronef, tel qu'un turboreacteur

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2061396A (en) * 1979-10-24 1981-05-13 Rolls Royce Turbine blade tip clearance control
US6742987B2 (en) 2002-07-16 2004-06-01 General Electric Company Cradle mounted turbine nozzle
US20080008584A1 (en) * 2006-07-06 2008-01-10 Siemens Power Generation, Inc. Cantilevered framework support for turbine vane
EP1908924A2 (fr) * 2006-10-03 2008-04-09 Rolls-Royce plc Montage d'aubage fixe pour moteur à turbine à gaz

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2061396A (en) * 1979-10-24 1981-05-13 Rolls Royce Turbine blade tip clearance control
US6742987B2 (en) 2002-07-16 2004-06-01 General Electric Company Cradle mounted turbine nozzle
US20080008584A1 (en) * 2006-07-06 2008-01-10 Siemens Power Generation, Inc. Cantilevered framework support for turbine vane
US7762766B2 (en) 2006-07-06 2010-07-27 Siemens Energy, Inc. Cantilevered framework support for turbine vane
EP1908924A2 (fr) * 2006-10-03 2008-04-09 Rolls-Royce plc Montage d'aubage fixe pour moteur à turbine à gaz
US8356981B2 (en) 2006-10-03 2013-01-22 Rolls-Royce Plc Gas turbine engine vane arrangement

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3082555A1 (fr) * 2018-06-18 2019-12-20 Safran Aircraft Engines Roue d'aubes redresseuses d'un moteur d'aeronef, tel qu'un turboreacteur

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