EP2211023A1 - Distributeur pour turbomachine avec structure support d'aubes directrices segmentée - Google Patents

Distributeur pour turbomachine avec structure support d'aubes directrices segmentée Download PDF

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Publication number
EP2211023A1
EP2211023A1 EP09000797A EP09000797A EP2211023A1 EP 2211023 A1 EP2211023 A1 EP 2211023A1 EP 09000797 A EP09000797 A EP 09000797A EP 09000797 A EP09000797 A EP 09000797A EP 2211023 A1 EP2211023 A1 EP 2211023A1
Authority
EP
European Patent Office
Prior art keywords
guide
vane
blade
guide vane
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Withdrawn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
EP09000797A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
Giuseppe Gaio
Holger Dr. Grote
Christian Lerner
Mirko Milazar
Mathias Stutt
Thomas-Dieter Tenrahm
Bernd Vonnemann
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.)
Siemens AG
Original Assignee
Siemens 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 Siemens AG filed Critical Siemens AG
Priority to EP09000797A priority Critical patent/EP2211023A1/fr
Priority to US13/145,353 priority patent/US9238976B2/en
Priority to JP2011545707A priority patent/JP5357270B2/ja
Priority to EP10700394.9A priority patent/EP2379846B1/fr
Priority to PCT/EP2010/050024 priority patent/WO2010084028A1/fr
Priority to CN201080004513.XA priority patent/CN102282340B/zh
Publication of EP2211023A1 publication Critical patent/EP2211023A1/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Withdrawn legal-status Critical Current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
    • 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
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/80—Repairing, retrofitting or upgrading methods

Definitions

  • the invention relates to a guide vane system, in particular for a gas turbine, with a number of vane rows and a vane carrier. It further relates to a gas turbine with such a guide vane system.
  • Gas turbines are used in many areas to drive generators or work machines.
  • the energy content of a fuel is used to generate a rotational movement of a turbine shaft.
  • the fuel is burned in a combustion chamber, compressed air being supplied by an air compressor.
  • the working medium produced in the combustion chamber by the combustion of the fuel, under high pressure and at high temperature, is guided via a turbine unit arranged downstream of the combustion chamber, where it relaxes to perform work.
  • a number of rotor blades which are usually combined into blade groups or rows of blades, are arranged thereon and drive the turbine shaft via a momentum transfer from the working medium.
  • guide vanes are also usually arranged between adjacent rotor blade rows and connected to the turbine housing, which are combined into rows of guide blades. These are attached via a blade root to a usually hollow cylindrical or hollow cone-shaped guide blade carrier and fastened on its side facing the turbine axis via a blade head on a common inner ring of the respective guide blade row. In stationary gas turbines, this often consists of an upper and a lower half, which are connected to each other via flanges.
  • the invention is therefore based on the object, a guide vane system, in particular for a gas turbine to specify, which allows a particularly simple exchange of vanes while maintaining a particularly high efficiency and thus designed for a particularly short repair time.
  • the guide vane carrier is segmented in the region of a guide vane row, wherein a segment extends over the entire radial extent extends the guide vane and the connection of the respective segment with the rest of the vane carrier is releasable.
  • the invention is based on the consideration that a shortened repair time would be possible by a particularly simple interchangeability of the vanes, if their assembly and disassembly could be simplified.
  • the guide vanes are within the guide vane carrier, which consists of an upper and a lower solid casting in stationary gas turbines, and thus also has to be disassembled to replace the vanes.
  • the guide vane carrier should therefore be segmented in the region of a guide vane row. This makes it possible to reach the area under the segments of the vane carrier only by uncovering individual segments.
  • a segment In order to ensure accessibility of the guide vanes, a segment should extend over the entire radial extent of the guide vane carrier and the connection of the respective segment to the rest of the vane carrier be solvable. Thus, for a repair or replacement of a vane no longer the entire vane carrier must be revealed, but it is only the connection of the respective segment with the rest of the vane support dissolved and - as the segment extends over the entire radial extent of the vane support - allows removal of the each segment directly reaching the vanes and replacing them.
  • the first turbine stage vane ie the vane closest to the combustor
  • the guide vane carrier should therefore advantageously be segmented in the region of the guide vane row closest to a combustion chamber of the gas turbine.
  • a plurality of segments should be provided, which extend in their entirety over the entire circumferential direction of the vane support.
  • an exchange of vanes by the respective segment is taken over the relevant vane.
  • the exact geometric design of the segmentation should be adapted in a meaningful way to the handling of the machine.
  • the respective connection is a screw connection and / or a tongue and groove connection.
  • the guide blade fixation should be provided in a meaningful way such that an undisturbed disassembly of an arbitrarily circumferentially located segment is ensured is, so that depending on the position of the blade to be replaced only the affected segment must be dismantled.
  • a guide vane of the respective vane row is releasably connected to the rest of the vane carrier.
  • the vane of the respective vane row is advantageously releasably connected on its side facing the turbine axis with an inner ring in the radial direction.
  • a radial distance of the vane is possible. This allows a particularly simple exchange.
  • the fixation of the guide vane on the inner ring is designed as a simple connector.
  • the respective vane advantageously comprises a spring which can be inserted in the radial direction into a groove of the inner ring.
  • the guide vanes of a row of guide vanes were fixed to the inner ring via a connection secured with pins, so that the entire inner ring had to be removed for removal and subsequently the guide vanes could be removed.
  • a releasable connection for example in the manner of a simple plug connection of the guide vanes with the inner ring should therefore the inner ring with the Brennschnabe, ie a fixed to the combustion chamber and thus the static part of the gas turbine component are fixed.
  • the inner ring is advantageously connected to a combustion chamber hub. This can be done for example by a fixation by welding, clamping o. ⁇ .
  • the inner ring can also be manufactured directly as part of the fuel chamber hub.
  • such a guide vane system is used in a gas turbine.
  • an outer housing of the gas turbine thereby advantageously comprises a manhole, through which easy access to the segments of the guide vane support for the assembly personnel is possible.
  • a gas and steam turbine power plant comprises such a gas turbine.
  • the advantages achieved by the invention are in particular that the segmentation of the guide vane carrier in the region of a guide vane row into segments which extend over the entire radial extent of the vane carrier and their connection with the rest of the vane carrier is solvable, a particularly simple exchange of vanes of a row of vanes is possible because not the entire turbine has to be revealed in such an exchange.
  • the cost of replacing the blades is significantly reduced and the required period of silence of the gas turbine can be significantly reduced.
  • Such a simplified replacement, in particular the first vane stage directly after the combustion chamber also allows an increase in the outlet temperature in conjunction with an increase in the efficiency of the gas turbine, since the simplified exchange option of the vanes on their durability less consideration must be taken. In this case, variable exchange concepts are conceivable during operation. Furthermore, such a design enables a comparatively faster test of new prototypes of guide vanes, for example with novel coatings or new cooling concepts, through the simplified exchange in research and development.
  • FIG. 1 shows a guide blade system 1 in sections in the region of the first two in hot gas direction to a combustion chamber 2 following vanes.
  • the illustration shows a half section through the upper half of a conically shaped guide blade carrier 4 and the respective vanes 6 of the first turbine stage and guide vanes 8 of the second turbine stage arranged in the vertex of the guide vane ring.
  • the vanes 6, 8 each comprise a blade root 10, 12 and a blade head 14, 16, via which their attachment to the other components takes place.
  • the guide vanes 6, 8 of the first and second turbine stage are fastened with their blade roots 10, 12 on the guide blade carrier 4 and at their respective blade heads 14, 16 fixed to inner rings 18, 20.
  • both the inner ring 20 and the vane support 4 include a plurality of cooling channels, not shown, which provide a cooling air supply to the guide vane 4, the vanes 6, 8 and the inner ring 20 to sufficiently cool these components due to the high hot gas temperatures.
  • the guide vane carrier 4 is segmented in the region of the first vane row.
  • the vane support 4 thus comprises a number of segments 24 and a remaining vane support 26 which is not segmented.
  • the segments 24 and the remaining vane support 25 are detachably connected.
  • the connection is effected via a hooking by means of grooves 28 and springs 30 introduced into the segments 24 and the remaining guide blade carrier 26.
  • Such a connection of the segments 24 is also provided with the combustion chamber wall 32.
  • the rest of the vane carrier 26 can also be understood as meaning an upper and / or a lower half of a ring-shaped vane carrier which is annular in cross-section, as is already known in stationary gas turbines.
  • at least two segments 24 are provided on each half of the remainder of the vane carrier.
  • more segments 24 are always provided in the axial section of the first row of guide vanes for the circumference than remaining guide blade carriers 26 in the adjoining axial section.
  • the vanes 6 of the first turbine stage can be reached without completely revealing the entire turbine from the outside.
  • the vane 6 of the first turbine stage is releasably secured via the blade root 10 by means of a fastening device 34 on the rest of the vane carrier 26.
  • this connection can be solved and the vane 6 is radially accessible from the outside.
  • the blade head 14 of the guide vane 6 of the first turbine stage in this case comprises a spring 36 which is inserted in a groove 38 of the inner ring 18.
  • the attachment to the inner ring 18 is thus designed only as a plug-in connection, so that the guide blade 6 can be easily removed to the outside after loosening the fastening device 34.
  • FIG. 2 also shows the vane system 1 as in FIG. 1
  • the detachable connection of the segment 24 with the rest of the guide blade carrier 26 is realized via a screw 40.
  • the entanglement of the segment 24 with the combustion chamber wall 32 via grooves 28 and springs 30 is unchanged.
  • Such a connection with a screw 40 may be desirable depending on strength or geometric requirements in the guide vane system 1.
  • FIG. 3 now shows a perpendicular to the turbine axis section through the vane system 1 at the height of the segments 24.
  • a total of twelve segments 24 are provided, which are connected via flanges 52, for example with a screw connection.
  • flanges 52 for example with a screw connection.
  • the segmentation can also be done in other ways and adapted to the handling of the machine.
  • FIG. 4 shows the combustion chamber hub 54 of a gas turbine.
  • This includes a groove 56 into which the in FIG. 1 and 2 shown inner ring 18 is used.
  • a groove 58 is provided, in which a sealing plate for sealing the gap between the blade root 14 of the guide vane 6 of the first turbine stage and the combustion chamber hub 54 is provided.
  • FIG. 5 shows a known attachment of the Leitschaufelfußes 14 to the combustion chamber hub 54 of the gas turbine in detail.
  • the blade root 14 comprises a spring 36 which is inserted into a groove 38 of the inner ring 18.
  • the guide vane of the first turbine stage 6 is fixed by means of a pin 60.
  • the inner ring 18 is then inserted into the groove 56 of the combustion chamber hub 54.
  • the blade root 14 includes a groove 62 for receiving a sealing plate 64, which also lies in the groove 58 of the combustion chamber hub 54.
  • the spring 36 of the blade root 14 is no longer connected via a pin to the inner ring 18 in the groove 38, but is merely stuck to the inner ring 18. Instead, the inner ring 18 is secured to the combustion chamber 54 by a pin 66 or screw. As a result, the guide vanes 6 can also be removed individually, without disassembling the inner ring 18. A secure hold of the guide vanes 6 is still on the fastening device 34, as in FIG. 1 and 2 shown, ensured.
  • FIG. 8 shows a section perpendicular to the turbine axis through two adjacent vanes 6 of the first turbine stage, as is common in the prior art.
  • grooves 68 are introduced into the blade feet 10 and blade heads 14 on the surface facing the adjacent guide blade 6, into which sealing plates 70 are inserted which close the gaps between the blade roots 10 and blade heads 14.
  • these sealing plates 70 may be a hindrance in a radial removal of individual vanes 6.
  • a plurality of guide vanes 6 are first to be unlocked and displaced in the circumferential direction, so that a guide blade 6 comes out of engagement of the sealing plates 70 and can be radially expanded.
  • Such a guide blade system 1 described here is advantageously used in a gas turbine 101.
  • a gas turbine 101 as in FIG. 10 has a compressor 102 for combustion air, a combustion chamber 2 and a turbine unit 106 for driving the compressor 102 and a generator, not shown, or a working machine.
  • the turbine unit 106 and the compressor 102 are arranged on a common, also referred to as a turbine rotor turbine shaft 108, with which the generator or the working machine is connected, and to its turbine axis 109 is rotatably mounted.
  • the running in the manner of an annular combustion chamber 2 is equipped with a number of burners 110 for the combustion of a liquid or gaseous fuel.
  • the turbine unit 106 has a number of rotatable blades 112 connected to the turbine shaft 108.
  • the blades 112 are annularly disposed on the turbine shaft 108 and thus form a number of blade rows.
  • the turbine unit 106 includes a number of stationary vanes 6, 8, 114 which are also annularly attached to a vane support 4 of the turbine unit 106 to form rows of vanes.
  • the blades 112 serve to drive the turbine shaft 108 by momentum transfer from the turbine unit 106 flowing through the working medium M.
  • the vanes 6, 8, 114 serve to guide the flow of the working medium M between two seen in the flow direction of the working medium M consecutive blade rows or blade rings.
  • a successive pair of a ring of vanes 114 or a row of vanes and a ring of blades 112 or a blade row is also referred to as a turbine stage.
  • Each vane 114 has a blade root 118, which is arranged to fix the respective vane 114 on a vane support 4 of the turbine unit 106 as a wall element.
  • Each blade 112 is fixed to the turbine shaft 108 in a similar manner via a blade root 119.
  • each ring segment 121 is arranged on the guide blade carrier 4 of the turbine unit 106.
  • the outer surface of each ring segment 121 is in the radial direction from the outer end of the blades 112 lying opposite it spaced by a gap.
  • the arranged between adjacent rows of stator ring segments 121 serve in particular as cover that protect the inner housing in the guide blade carrier 4 or other housing-mounting components from thermal overload by the turbine 106 flowing through the hot working medium M.
  • the combustion chamber 2 is configured in the exemplary embodiment as a so-called annular combustion chamber, in which a plurality of burners 110 arranged around the turbine shaft 108 in the circumferential direction open into a common combustion chamber space.
  • the combustion chamber 2 is configured in its entirety as an annular structure, which is positioned around the turbine shaft 108 around.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP09000797A 2009-01-21 2009-01-21 Distributeur pour turbomachine avec structure support d'aubes directrices segmentée Withdrawn EP2211023A1 (fr)

Priority Applications (6)

Application Number Priority Date Filing Date Title
EP09000797A EP2211023A1 (fr) 2009-01-21 2009-01-21 Distributeur pour turbomachine avec structure support d'aubes directrices segmentée
US13/145,353 US9238976B2 (en) 2009-01-21 2010-01-05 Guide vane system for a turbomachine having segmented guide vane carriers
JP2011545707A JP5357270B2 (ja) 2009-01-21 2010-01-05 分割ガイドベーンキャリアを有するターボ機械用のガイドベーンシステム
EP10700394.9A EP2379846B1 (fr) 2009-01-21 2010-01-05 Structure de support d'aubes directrices pour turbomachine
PCT/EP2010/050024 WO2010084028A1 (fr) 2009-01-21 2010-01-05 Système d'aubes directrices pour une turbomachine comportant un support d'aubes directrices segmenté
CN201080004513.XA CN102282340B (zh) 2009-01-21 2010-01-05 具有分段的导向叶片外圈的涡轮机导向叶片***

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP09000797A EP2211023A1 (fr) 2009-01-21 2009-01-21 Distributeur pour turbomachine avec structure support d'aubes directrices segmentée

Publications (1)

Publication Number Publication Date
EP2211023A1 true EP2211023A1 (fr) 2010-07-28

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ID=40942519

Family Applications (2)

Application Number Title Priority Date Filing Date
EP09000797A Withdrawn EP2211023A1 (fr) 2009-01-21 2009-01-21 Distributeur pour turbomachine avec structure support d'aubes directrices segmentée
EP10700394.9A Active EP2379846B1 (fr) 2009-01-21 2010-01-05 Structure de support d'aubes directrices pour turbomachine

Family Applications After (1)

Application Number Title Priority Date Filing Date
EP10700394.9A Active EP2379846B1 (fr) 2009-01-21 2010-01-05 Structure de support d'aubes directrices pour turbomachine

Country Status (5)

Country Link
US (1) US9238976B2 (fr)
EP (2) EP2211023A1 (fr)
JP (1) JP5357270B2 (fr)
CN (1) CN102282340B (fr)
WO (1) WO2010084028A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2442033A1 (fr) * 2010-10-12 2012-04-18 Siemens Aktiengesellschaft Segment d'accrochage de chambre de combustion et coque extérieure de chambre de combustion
US20170226887A1 (en) * 2016-02-05 2017-08-10 MTU Aero Engines AG Guide vane system for a turbomachine

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EP2644833A1 (fr) * 2012-03-26 2013-10-02 Alstom Technology Ltd Anneau de support
EP2692995B1 (fr) * 2012-07-30 2017-09-20 Ansaldo Energia IP UK Limited Moteur à turbine à gaz stationnaire et procédé pour effectuer les travaux de maintenance
FR3008912B1 (fr) * 2013-07-29 2017-12-15 Snecma Carter de turbomachine et procede de fabrication
US10072516B2 (en) 2014-09-24 2018-09-11 United Technologies Corporation Clamped vane arc segment having load-transmitting features
DE102015224988A1 (de) * 2015-12-11 2017-06-14 Rolls-Royce Deutschland Ltd & Co Kg Verfahren zur Montage einer Brennkammer eines Gasturbinentriebwerks
JP6763157B2 (ja) * 2016-03-11 2020-09-30 株式会社Ihi タービンノズル
US20180106155A1 (en) * 2016-10-13 2018-04-19 Siemens Energy, Inc. Transition duct formed of a plurality of segments
DE102017204953A1 (de) 2017-03-23 2018-09-27 MTU Aero Engines AG Strömungsmaschine, Verfahren und Leitschaufelreihensystem

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GB1427915A (en) * 1967-12-19 1976-03-10 Gen Motors Corp Gas turbine cooling
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GB2240822A (en) * 1990-01-16 1991-08-14 Gen Electric Improved arrangement for sealing gaps between segments of, e.g. turbine nozzles and shrouds
US20050132707A1 (en) * 2001-11-20 2005-06-23 Andreas Gebhardt Gas turbo set
US20060032236A1 (en) * 2004-06-17 2006-02-16 Snecma Moteurs Mounting a high pressure turbine nozzle in leaktight manner to one end of a combustion chamber in a gas turbine
EP1840337A1 (fr) * 2006-03-31 2007-10-03 Siemens Aktiengesellschaft Joint de rainure et languette entre deux composants de turbine
WO2008012195A1 (fr) * 2006-07-24 2008-01-31 Siemens Aktiengesellschaft Procédé pour dévisser une moitié annulaire d'un distributeur de forme globale annulaire hors d'une moitié inférieure de boîtier d'une turbomachine stationnaire à écoulement axial, dispositif de montage, assemblage de dispositif de montage et demi-secteur annulaire auxiliaire

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Publication number Priority date Publication date Assignee Title
US3300180A (en) * 1964-11-17 1967-01-24 Worthington Corp Segmented diaphragm assembly
GB1427915A (en) * 1967-12-19 1976-03-10 Gen Motors Corp Gas turbine cooling
US4083648A (en) * 1975-08-01 1978-04-11 United Technologies Corporation Gas turbine construction
GB2240822A (en) * 1990-01-16 1991-08-14 Gen Electric Improved arrangement for sealing gaps between segments of, e.g. turbine nozzles and shrouds
US20050132707A1 (en) * 2001-11-20 2005-06-23 Andreas Gebhardt Gas turbo set
US20060032236A1 (en) * 2004-06-17 2006-02-16 Snecma Moteurs Mounting a high pressure turbine nozzle in leaktight manner to one end of a combustion chamber in a gas turbine
EP1840337A1 (fr) * 2006-03-31 2007-10-03 Siemens Aktiengesellschaft Joint de rainure et languette entre deux composants de turbine
WO2008012195A1 (fr) * 2006-07-24 2008-01-31 Siemens Aktiengesellschaft Procédé pour dévisser une moitié annulaire d'un distributeur de forme globale annulaire hors d'une moitié inférieure de boîtier d'une turbomachine stationnaire à écoulement axial, dispositif de montage, assemblage de dispositif de montage et demi-secteur annulaire auxiliaire

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2442033A1 (fr) * 2010-10-12 2012-04-18 Siemens Aktiengesellschaft Segment d'accrochage de chambre de combustion et coque extérieure de chambre de combustion
US20170226887A1 (en) * 2016-02-05 2017-08-10 MTU Aero Engines AG Guide vane system for a turbomachine
US10450888B2 (en) * 2016-02-05 2019-10-22 MTU Aero Engines AG Guide vane system for a turbomachine

Also Published As

Publication number Publication date
WO2010084028A1 (fr) 2010-07-29
EP2379846B1 (fr) 2019-11-06
US20120039716A1 (en) 2012-02-16
CN102282340B (zh) 2016-01-20
CN102282340A (zh) 2011-12-14
JP2012515869A (ja) 2012-07-12
US9238976B2 (en) 2016-01-19
EP2379846A1 (fr) 2011-10-26
JP5357270B2 (ja) 2013-12-04

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