EP2379846A1 - Système d'aubes directrices pour une turbomachine comportant un support d'aubes directrices segmenté - Google Patents

Système d'aubes directrices pour une turbomachine comportant un support d'aubes directrices segmenté

Info

Publication number
EP2379846A1
EP2379846A1 EP10700394A EP10700394A EP2379846A1 EP 2379846 A1 EP2379846 A1 EP 2379846A1 EP 10700394 A EP10700394 A EP 10700394A EP 10700394 A EP10700394 A EP 10700394A EP 2379846 A1 EP2379846 A1 EP 2379846A1
Authority
EP
European Patent Office
Prior art keywords
turbine
vane
segments
guide
gas turbine
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.)
Granted
Application number
EP10700394A
Other languages
German (de)
English (en)
Other versions
EP2379846B1 (fr
Inventor
Fathi Ahmad
Giuseppe Gaio
Holger 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 EP10700394.9A priority Critical patent/EP2379846B1/fr
Publication of EP2379846A1 publication Critical patent/EP2379846A1/fr
Application granted granted Critical
Publication of EP2379846B1 publication Critical patent/EP2379846B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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
    • 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

  • 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.
  • For flow guidance of the working medium are also usually connected between adjacent rows of blades with the turbine housing, too
  • the invention is therefore based on the object of specifying a turbine guide vane carrier, in particular for a gas turbine, which, while maintaining a particularly high efficiency, enables a particularly simple exchange of individual vanes and is therefore designed for a particularly short repair duration.
  • the vane of the first turbine stage ie the vane, which is the combustor chamber is closest to these highest temperatures and subject to the greatest wear. Accordingly, premature replacement as a result of damage caused by clogging of the cooling air bores (eg by inwardly oxidizing cooling air bores) is to be expected in particular in the case of this turbine guide vane.
  • the vane support should therefore advantageously be segmented multiple times in the section of the vane row closest to a combustion chamber of the gas turbine.
  • the upstream side portion of the turbine vane carrier should have more segments than the remaining portion of the turbine vane carrier.
  • the guide blade fixation of a gas turbine should be provided in a meaningful manner such that undisturbed disassembly of an arbitrarily circumferentially located segment is ensured, so that, depending on the Position of the blade to be replaced only the affected, radially further outwardly disposed segment must be dismantled.
  • the guide vane of the respective vane row with one of the segments of the remaining portion releasably connected. Thereby, after removal of the affected segment, the vane can be disassembled by releasing the connection with the segment of the section.
  • the segments arranged in the inflow-side section thus do not serve for fastening guide vanes, but only for producing or maintaining the integrity of the gas turbine and possibly for separating spaces for cooling air with different pressures and / or temperatures.
  • such a turbine guide vane carrier is used in a gas turbine.
  • an outer housing of the gas turbine advantageously comprises a manhole through which easy access to the segments of the vane support for the installation personnel is possible.
  • FIG. 9 shows a section through two adjacent vanes perpendicular to the turbine axis with clamping elements fixed sealing elements, and 10 shows a half section through a gas turbine.
  • 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 stages are fastened to the guide blade carrier 1 with their blade roots 10, 12 and fixed to inner rings 18, 20 at their respective blade heads 14, 16.
  • both the inner ring 20 and the vane support 1 include a plurality of cooling systems 22, which provide a cooling air supply to the vane support 1, the vanes 6, 8 and the inner ring 22 to sufficiently cool these components due to the high hot gas temperatures.
  • the guide vane carrier 1 is segmented in many cases in the region of the first vane row.
  • the guide blade carrier 1 comprises a number (in this case 12 pieces, see FIG. 3) of segments 24 in an inflow-side portion 23 and a guide-blade carrier 1 segmented in only one half 25 in a remaining portion 25. All segments 24, 26 are detachably connected to one another.
  • the rest of the vane carrier is understood to be an upper and a lower half of a ring-shaped vane carrier which is annular in cross-section, as is already known in stationary gas turbines. In this case, two
  • FIG. 3 shows a section perpendicular to the turbine axis through the guide vane carrier 1 at the level 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.
  • 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 a groove 62 for receiving a sealing plate 64, which is also located in the groove 58 of the combustion chamber hub 54.
  • a plurality of guide vanes 6 are first to be unlocked and displaced in the circumferential direction, so that a guide Schaufei 6 comes out of engagement of the sealing plates 70 and can be expanded radially.
  • a gas turbine 101 has a compressor 102 for combustion air, a combustion chamber 2 and a turbine unit 106 for driving the compressor 102 and a generator or a working machine (not shown).
  • the turbine unit 106 and the compressor 102 are arranged on a common turbine shaft 108, also referred to as a turbine rotor, to which the generator or the working machine is also connected, and which is rotatably mounted about its turbine axis 109.
  • 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 attached in a donut-like manner to a vane support 1 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 guide vanes and of a ring of rotor blades 112 or a row of revolutions 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 1 of the turbine unit 106 as a wall element.
  • Each rotor blade 112 is fastened 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 1 of the turbine unit 106.
  • the outer surface of each ring segment 121 is spaced apart in the radial direction from the outer end of the blades 112 lying opposite it by a gap. The between adjacent
  • 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.

Landscapes

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

Abstract

L'invention concerne un support d'aubes directrices de turbine (1) notamment destiné à une turbine à gaz (101), comportant une pluralité de rangées d'aubes directrices et un support d'aubes directrices (1). L'invention vise à mettre en oeuvre un tel support d'aubes directrices de turbine (1) permettant un échange particulièrement simple des aubes directrices tout en conservant un rendement élevé, et étant donc adapté à une durée de réparation particulièrement courte. A cet effet, le support d'aubes directrices de turbine (1) comporte une pluralité de segments (24, 26) s'étendant sur toute l'extension radiale du support d'aubes directrices de turbine (1); la liaison des segments restants (26) est amovible; et le support d'aubes directrices de turbine (1) comporte, le long de son extension axiale, au moins deux sections (23, 25) connectées l'une à l'autre, comprenant un nombre de segments (24, 26) différent.
EP10700394.9A 2009-01-21 2010-01-05 Structure de support d'aubes directrices pour turbomachine Active EP2379846B1 (fr)

Priority Applications (1)

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

Applications Claiming Priority (3)

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
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

Publications (2)

Publication Number Publication Date
EP2379846A1 true EP2379846A1 (fr) 2011-10-26
EP2379846B1 EP2379846B1 (fr) 2019-11-06

Family

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 Before (1)

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

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)

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EP2442033A1 (fr) * 2010-10-12 2012-04-18 Siemens Aktiengesellschaft Segment d'accrochage de chambre de combustion et coque extérieure de chambre de combustion
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
DE102016201766A1 (de) * 2016-02-05 2017-08-10 MTU Aero Engines AG Leitschaufelsystem für eine Strömungsmaschine
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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Also Published As

Publication number Publication date
JP2012515869A (ja) 2012-07-12
EP2211023A1 (fr) 2010-07-28
WO2010084028A1 (fr) 2010-07-29
CN102282340A (zh) 2011-12-14
CN102282340B (zh) 2016-01-20
US20120039716A1 (en) 2012-02-16
US9238976B2 (en) 2016-01-19
EP2379846B1 (fr) 2019-11-06
JP5357270B2 (ja) 2013-12-04

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