EP2282014A1 - Section de canal d'écoulement annulaire pour une turbomachine - Google Patents

Section de canal d'écoulement annulaire pour une turbomachine Download PDF

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Publication number
EP2282014A1
EP2282014A1 EP09008227A EP09008227A EP2282014A1 EP 2282014 A1 EP2282014 A1 EP 2282014A1 EP 09008227 A EP09008227 A EP 09008227A EP 09008227 A EP09008227 A EP 09008227A EP 2282014 A1 EP2282014 A1 EP 2282014A1
Authority
EP
European Patent Office
Prior art keywords
flow channel
channel section
platforms
shielding
platform
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Withdrawn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
EP09008227A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
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 EP09008227A priority Critical patent/EP2282014A1/fr
Priority to EP10725431A priority patent/EP2446119A1/fr
Priority to CN2010800282474A priority patent/CN102803658A/zh
Priority to PCT/EP2010/058352 priority patent/WO2010149528A1/fr
Priority to US13/379,530 priority patent/US20120100008A1/en
Priority to JP2012516636A priority patent/JP5443600B2/ja
Publication of EP2282014A1 publication Critical patent/EP2282014A1/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Withdrawn legal-status Critical Current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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • F01D11/006—Sealing the gap between rotor blades or blades and rotor
    • F01D11/008—Sealing the gap between rotor blades or blades and rotor by spacer elements between the blades, e.g. independent interblade platforms
    • 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
    • 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
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms
    • 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
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid

Definitions

  • the invention relates to an annular flow channel section for a turbomachine, comprising a vane ring having a number of circumferentially juxtaposed vanes each comprising a blade root, a platform and a radiantly projecting into the flow channel airfoil, wherein the flow channel is platform-side limited by shielding, each sit between two immediately adjacent blades.
  • annular flow channel section is for example from the EP 1 219 787 B1 known.
  • the patent discloses a ring of cast guide vanes of an axial flow turbine in which the vanes have an aerodynamically curved airfoil at each of whose radially outer (foot-side) and inner (head-side) ends platforms are provided. Installed in the turbine, the platforms are covered by ceramic heat shields.
  • the heat shields are configured such that they each cover one half of the platform of two directly adjacent vanes in pairs. They thus essentially extend from the suction side wall of the blade profile of a first guide blade to the pressure side wall of the blade profile of a second guide blade.
  • the ceramic heat shield is connected via a spring fixed to the gas turbine blade, so that the former is attached interchangeable.
  • ceramic heat shields require a comparatively large wall thickness in order to be able to permanently and reliably withstand the temperatures of the hot gas occurring in a stationary gas turbine. If such ceramic heat shields on both the head-side and on the foot side Can be used platform of vanes, this leads to relatively large turbine vanes with correspondingly increased space requirements, which also increases the cost of production.
  • the object of the invention is therefore to provide an annular flow channel section for a turbomachine, which requires a comparatively small space requirement and, moreover, for a particularly long period of time the hot gas flowing in the flow channel section leads particularly reliably and safely without premature signs of wear occurring at the flow channel defining components ,
  • the object is achieved with an annular flow channel section for a turbomachine, in which the shielding elements are arranged under gap formation on the platforms and in the platform impingement cooling openings are provided for impingement cooling of the shielding elements.
  • the invention is based on the finding that the platform integrally formed on the guide vanes can be protected from the hot gas and its corrosive and thermal influences even when the shielding element is not made of a ceramic. In this case, the shielding is then sufficient to cool. According to the invention, it is provided that an impingement cooling of the shielding element is used for cooling. By cooling the shielding this can be configured thin-walled than in the prior art. The comparatively thin-walled design of the shielding element is space-saving and also less expensive. The airfoil of respective vanes can thereby be made shorter in its span, without reducing the flow area of the annular flow channel portion, compared with the known from the prior art flow channel section.
  • the shielding elements preferably each have a base plate which delimits the flow channel and is made of a metallic material which is manufactured separately from the guide vanes. By cooling the shielding can be made of metallic materials. In addition, the entire shielding element is manufactured separately from the guide vanes. This has the advantage that in the event of wear and tear on the shielding only this is to be replaced and not the complete vane, as in non-shielded vane platforms.
  • the shielding element is made of a metallic material having good insulating properties.
  • the wall thickness of the base plate is less than the wall thickness of the covered by the shielding platform.
  • the thinner the shielding element the better this can be cooled by the impingement cooling.
  • a flow channel section that is compact in terms of space can be specified, which reduces the manufacturing and material costs for such a flow channel section.
  • transverse wall sections are provided at the edges of the base plate, which are connectable to lateral walls of the platforms. This makes it possible to accomplish a convenient attachment of the shielding to the vane.
  • each shield member extends over a gap bounded by the platforms of two immediately adjacent vanes. This allows a low-loss guidance of the hot gas in the flow channel, even in the event that, due to thermally induced strains, an offset of adjacent platforms occurs.
  • the shielding element has a protective layer on the flow channel side, in particular a heat-insulating protective layer.
  • FIG. 1 shows the cross section through the blades 14 of two vanes 10 of an annular flow channel portion 12 of a hot gas axially flowed through by a turbomachine, such as gas turbine.
  • the flow channel section 12 essentially comprises a vane ring with a plurality of circumferentially juxtaposed vanes 10.
  • FIG. 1 shows only two of the vanes 10 are shown.
  • the vanes 10 are attached in a conventional manner to a guide vane.
  • the representation is in FIG. 1 chosen so that the blades 14 are shown in cross section and thus a plan view of the platforms 16 of the vanes 10 takes place. Between a suction side airfoil wall 18 of in FIG.
  • the vane ring between each pair of immediately adjacent airfoils 14 each have such a shielding element 22, wherein adjacent shielding elements 22 on the one hand upstream of a leading edge 21 of the airfoil 14 and downstream of a trailing edge 23 of the airfoil 14 with the smallest possible gap abut each other.
  • baffle cooling openings 24 are arranged, for example, grid-shaped.
  • the section along the section line II-II through the vane 10 and the shield 22 shows FIG. 2 , In FIG. 2 are closed FIG. 1 identical features provided with identical reference numerals.
  • the shielding element 22 is arranged with gap formation on the platform 16 on the hot gas side, wherein in the platform 16 to the surface of which, for example, oblique impingement cooling openings 24 are provided.
  • a coolant K is supplied during operation of the turbomachine, which emerge through the impingement cooling openings 24 from the rear space 28 and can enter into the gap between shield 22 and platform 16 like a jet. When the impact cooling jets strike, they cool the shielding element 22, so that, despite the hot gas flowing through the flow channel 26, it has a sufficient service life.
  • Shielding element 22 shown in cross-section is metallic and essentially comprises a base plate 30 which extends parallel to the channel-side platform surface. At the two opposite edges of the base plate 30 laterally transversely to the base plate 30 projecting wall portions 32 are provided, which surround respective side walls of the platform 16 like a clamp.
  • the Wall thickness of the base plate 30 is substantially lower than the wall thickness of the platform 16 in the region of the impact cooling openings 24.
  • the shielding element 22 For fastening the shielding element 22 to the guide blade 10 or to the platform 16, this can be screwed, for example, as indicated by the dot-dash line. Other types of attachment such. As well as a jamming, in particular positive clamping of the shielding member 22 to the platform 16 is also conceivable. If necessary, the shielding member 22 may have on its surface, which is exposed to the hot gas, a thermal heat-insulating layer in order to further increase its thermal resistance.
  • the coolant K flowing into the gap between the shielding element 22 and the platform surface flows after the impingement cooling at that gap 36 (FIG. FIG. 1 ), which is provided between the shielding member 22 and the suction-side airfoil wall 18 and pressure-side airfoil wall 20, respectively.
  • platform 16 and the shielding member 22 disposed above it may be both a foot-side platform and a head-side platform of vanes 10, provided that the guide vanes 10 used in the annular flow channel section 12 at both opposite ends of the blade 14 transverse to the blade 14th have extending platforms 16.
  • the invention can also be applied to only one of the two platforms 16 of such a vane 10.
  • the invention provides an annular flow channel section 12 for a turbomachine comprising a vane ring which has a number of circumferentially juxtaposed vanes 10, each comprising a platform 16 and a radiant manner into the flow channel 26 projecting airfoil 14, wherein the flow channel 26 platform side is limited by shielding 22, which are each arranged between two immediately adjacent airfoils 14, wherein the formation of a particularly space-saving flow channel section 12, the shielding 22 arranged under gap formation on the platforms 16 and in the platform 16th Impact cooling holes 24 are provided.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP09008227A 2009-06-23 2009-06-23 Section de canal d'écoulement annulaire pour une turbomachine Withdrawn EP2282014A1 (fr)

Priority Applications (6)

Application Number Priority Date Filing Date Title
EP09008227A EP2282014A1 (fr) 2009-06-23 2009-06-23 Section de canal d'écoulement annulaire pour une turbomachine
EP10725431A EP2446119A1 (fr) 2009-06-23 2010-06-15 Segment de canal d'écoulement de forme annulaire pour une turbomachine
CN2010800282474A CN102803658A (zh) 2009-06-23 2010-06-15 用于涡轮机的环形的流动通道区段
PCT/EP2010/058352 WO2010149528A1 (fr) 2009-06-23 2010-06-15 Segment de canal d'écoulement de forme annulaire pour une turbomachine
US13/379,530 US20120100008A1 (en) 2009-06-23 2010-06-15 Annular flow channel section for a turbomachine
JP2012516636A JP5443600B2 (ja) 2009-06-23 2010-06-15 ターボ機械のための環状流路

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP09008227A EP2282014A1 (fr) 2009-06-23 2009-06-23 Section de canal d'écoulement annulaire pour une turbomachine

Publications (1)

Publication Number Publication Date
EP2282014A1 true EP2282014A1 (fr) 2011-02-09

Family

ID=41351921

Family Applications (2)

Application Number Title Priority Date Filing Date
EP09008227A Withdrawn EP2282014A1 (fr) 2009-06-23 2009-06-23 Section de canal d'écoulement annulaire pour une turbomachine
EP10725431A Withdrawn EP2446119A1 (fr) 2009-06-23 2010-06-15 Segment de canal d'écoulement de forme annulaire pour une turbomachine

Family Applications After (1)

Application Number Title Priority Date Filing Date
EP10725431A Withdrawn EP2446119A1 (fr) 2009-06-23 2010-06-15 Segment de canal d'écoulement de forme annulaire pour une turbomachine

Country Status (5)

Country Link
US (1) US20120100008A1 (fr)
EP (2) EP2282014A1 (fr)
JP (1) JP5443600B2 (fr)
CN (1) CN102803658A (fr)
WO (1) WO2010149528A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2634373A1 (fr) * 2012-02-28 2013-09-04 Siemens Aktiengesellschaft Agencement pour turbomachine
EP2540971B1 (fr) * 2011-06-27 2019-04-03 General Electric Company Procédé de fabrication d'un passage de refroidissement de plate-forme dans une aube de rotor de turbine, et aube de rotor de turbine associé

Families Citing this family (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2014204608A1 (fr) * 2013-06-17 2014-12-24 United Technologies Corporation Pale de turbine ayant un patin de plateforme
ITCO20130051A1 (it) * 2013-10-23 2015-04-24 Nuovo Pignone Srl Metodo per la produzione di uno stadio di una turbina a vapore
JP6366180B2 (ja) 2014-09-26 2018-08-01 三菱日立パワーシステムズ株式会社 シール構造
US10598029B2 (en) * 2016-11-17 2020-03-24 United Technologies Corporation Airfoil with panel and side edge cooling
US20200182085A1 (en) * 2018-12-07 2020-06-11 United Technoligies Corporation Impingement cooling of components
CN112943378B (zh) * 2021-02-04 2022-06-28 大连理工大学 一种涡轮叶片枝网式冷却结构

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US5820336A (en) * 1994-11-11 1998-10-13 Mitsubishi Heavy Industries, Ltd. Gas turbine stationary blade unit
WO1999054597A1 (fr) * 1998-04-21 1999-10-28 Siemens Aktiengesellschaft Pale de turbine
WO1999060253A1 (fr) * 1998-05-18 1999-11-25 Siemens Aktiengesellschaft Plate-forme d'aube de turbine a refroidissement
EP1557534A1 (fr) * 2004-01-20 2005-07-27 Siemens Aktiengesellschaft Aube de turbine à gaz et turbine à gaz avec une telle aube
EP1557535A1 (fr) * 2004-01-20 2005-07-27 Siemens Aktiengesellschaft Aube de turbine à gaz et turbine à gaz avec une telle aube
EP1219787B1 (fr) 2000-12-27 2005-12-21 Siemens Aktiengesellschaft Aube de turbine à gaz et turbine à gaz
US20070237630A1 (en) * 2006-04-11 2007-10-11 Siemens Power Generation, Inc. Vane shroud through-flow platform cover

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BE755567A (fr) * 1969-12-01 1971-02-15 Gen Electric Structure d'aube fixe, pour moteur a turbines a gaz et arrangement de reglage de temperature associe
US4218178A (en) * 1978-03-31 1980-08-19 General Motors Corporation Turbine vane structure
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FR2758855B1 (fr) * 1997-01-30 1999-02-26 Snecma Systeme de ventilation des plates-formes des aubes mobiles
JP3546135B2 (ja) * 1998-02-23 2004-07-21 三菱重工業株式会社 ガスタービン動翼のプラットフォーム
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FR2810365B1 (fr) * 2000-06-15 2002-10-11 Snecma Moteurs Systeme de ventilation d'une paire de plates-formes d'aubes juxtaposees
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DE502005010381D1 (de) * 2005-04-28 2010-11-25 Siemens Ag Verfahren und Vorrichtung zur Einstellung eines Radialspaltes eines axial durchströmten Verdichters einer Strömungsmaschine
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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5820336A (en) * 1994-11-11 1998-10-13 Mitsubishi Heavy Industries, Ltd. Gas turbine stationary blade unit
WO1999054597A1 (fr) * 1998-04-21 1999-10-28 Siemens Aktiengesellschaft Pale de turbine
WO1999060253A1 (fr) * 1998-05-18 1999-11-25 Siemens Aktiengesellschaft Plate-forme d'aube de turbine a refroidissement
EP1219787B1 (fr) 2000-12-27 2005-12-21 Siemens Aktiengesellschaft Aube de turbine à gaz et turbine à gaz
EP1557534A1 (fr) * 2004-01-20 2005-07-27 Siemens Aktiengesellschaft Aube de turbine à gaz et turbine à gaz avec une telle aube
EP1557535A1 (fr) * 2004-01-20 2005-07-27 Siemens Aktiengesellschaft Aube de turbine à gaz et turbine à gaz avec une telle aube
US20070237630A1 (en) * 2006-04-11 2007-10-11 Siemens Power Generation, Inc. Vane shroud through-flow platform cover

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2540971B1 (fr) * 2011-06-27 2019-04-03 General Electric Company Procédé de fabrication d'un passage de refroidissement de plate-forme dans une aube de rotor de turbine, et aube de rotor de turbine associé
EP2634373A1 (fr) * 2012-02-28 2013-09-04 Siemens Aktiengesellschaft Agencement pour turbomachine
WO2013127833A1 (fr) * 2012-02-28 2013-09-06 Siemens Aktiengesellschaft Agencement pour une turbomachine
CN104136720A (zh) * 2012-02-28 2014-11-05 西门子公司 用于涡轮机器的装置
CN104136720B (zh) * 2012-02-28 2016-08-31 西门子公司 用于涡轮机器的装置
US9863271B2 (en) 2012-02-28 2018-01-09 Siemens Aktiengesellschaft Arrangement for a turbomachine

Also Published As

Publication number Publication date
EP2446119A1 (fr) 2012-05-02
JP2012530870A (ja) 2012-12-06
CN102803658A (zh) 2012-11-28
US20120100008A1 (en) 2012-04-26
WO2010149528A1 (fr) 2010-12-29
JP5443600B2 (ja) 2014-03-19

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