EP2446119A1 - Segment de canal d'écoulement de forme annulaire pour une turbomachine - Google Patents

Segment de canal d'écoulement de forme annulaire pour une turbomachine

Info

Publication number
EP2446119A1
EP2446119A1 EP10725431A EP10725431A EP2446119A1 EP 2446119 A1 EP2446119 A1 EP 2446119A1 EP 10725431 A EP10725431 A EP 10725431A EP 10725431 A EP10725431 A EP 10725431A EP 2446119 A1 EP2446119 A1 EP 2446119A1
Authority
EP
European Patent Office
Prior art keywords
flow channel
platform
shielding
channel section
platforms
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Withdrawn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
EP10725431A
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 EP10725431A priority Critical patent/EP2446119A1/fr
Publication of EP2446119A1 publication Critical patent/EP2446119A1/fr
Withdrawn legal-status Critical Current

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
    • 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 known, for example, from EP 1 219 787 B1.
  • 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 designed such that they each cover one half of the platform of two immediately 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.
  • An alternative construction for mounting such a cover is given in US 2007/0237630 Al.
  • ceramic heat shields require a comparatively large wall thickness in order to permanently and reliably protect the temperatures of the hot gas occurring in a stationary gas turbine. to be able to withstand. If such ceramic heat shields are used both on the head-side and on the foot-side platform of guide vanes, this leads to comparatively large turbine guide vanes with a correspondingly larger space requirement, which likewise increases the production costs.
  • EP 1 557 535 A1 discloses a modular turbine blade with two sheet metal shells which, in addition to the associated platform half, also covers the transition to the aerodynamically curved blade leaf.
  • the disadvantage here however, the sealing of the gap between abutting platform halves of adjacent turbine blades with the aid of a sealing element inserted in grooves.
  • a modified form of limiting the flow channel is shown in EP 1 557 534 A1.
  • 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 reliably and reliably conducts the hot gas flowing in the flow channel section for a particularly long period of time without premature failure at the components bordering the flow channel Wear and tear occur.
  • the object is achieved with an annular flow channel section for a turbomachine, in which the shielding elements are arranged with gap formation on the platforms and in the platform impingement cooling openings are provided for impingement cooling of the shield elements.
  • the invention is based on the finding that the platform halves 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.
  • each shielding element extends over a gap bounded by the platforms of two directly adjacent guide 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 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 velvet shielding made separately from the 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.
  • Wall thickness of the base plate less than the wall thickness of the shielding element covered platform.
  • 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.
  • the shielding element has a protective layer on the flow channel side, in particular a heat-insulating protective layer.
  • FIG. 1 shows a section through two of the airfoils of an annular flow channel section as a development ment with a arranged over the platforms of the vanes shielding and
  • FIG 2 shows the section according to section II-II through the platform of the vane and through the shielding.
  • FIG. 1 shows the cross section through the blades 14 of two guide vanes 10 of an annular Strömungskanalab- section 12 of an axially flowed through by a hot gas
  • the flow channel section 12 essentially comprises a guide vane ring with a plurality of guide vanes 10 which are lined up in the circumferential direction. Only two of the guide vanes 10 are shown in FIG. 1 of the guide vane ring which is widely known in the prior art.
  • the vanes 10 are attached in a conventional manner to a guide vane. The illustration is chosen in FIG. 1 so that the blades 14 are shown in cross-section and thus a plan view of the platforms 16 of the guide blades 10 takes place.
  • a shielding element 22 is arranged in a form-fitting manner. The shielding element 22 is in
  • baffle cooling openings 24 are arranged, for example, grid-shaped.
  • the section according to the section line II-II through the guide blade 10 and the shield element 22 is shown in FIG. 2.
  • identical features to FIG. 1 are provided with identical reference symbols.
  • the shielding element 22 is arranged with gap formation on the platform 16 on the hot gas side, wherein in the platform 16, for example, obliquely extending impact cooling openings 24 are provided to the surface thereof.
  • a coolant K which emerges from the rear space 28 through the impingement cooling openings 24 and can enter the gap between the shielding element 22 and the platform 16 in the manner of a jet, is supplied to the rear space 28 facing away from the flow channel 26.
  • 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.
  • the shielding element 22 shown in cross-section in FIG. 2 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 transverse to the base plate 30 projecting wall sections 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 smaller 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 on its surface, which is exposed to the hot gas, a thermal Have heat insulation layer to further increase its thermal resistance.
  • Impingement cooling at that gap 36 (FIG. 1), which is provided between the shielding element 22 and the suction-side airfoil wall 18 or pressure-side airfoil wall 20.
  • the platform 16 shown in FIG. 2 and the shielding element 22 arranged above it may be both a foot-side platform and a head-side platform of guide vanes 10, provided that the guide vanes 10 used in the annular flow channel section 12 at both opposite ends of the airfoil 14th have transverse to the blade 14 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 having a number of circumferentially juxtaposed vanes 10, each comprising a platform 16 and a radiant blade 14 projecting radially into the flow channel 26, the flow channel 26 platform side is delimited by shielding elements 22, which are each arranged between two immediately adjacent blades 14, wherein to form a particularly space-saving flow channel section 12, the shielding elements 22 are arranged to form gaps on the platforms 16 and in the platform 16 impact cooling openings 24 are provided.

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 canal d'écoulement (12) de forme annulaire pour une turbomachine, comportant une couronne d'aubes directrices qui présente un certain nombre d'aubes directrices (10) alignées les unes à coté des autres dans le sens périphérique et pourvues chacune d'un pied, d'une plate-forme (16) et d'une ailette (14) pénétrant en forme de rayon dans le canal d'écoulement (26), le canal d'écoulement (26) étant délimité, côté plate-forme, par des éléments de protection (22) disposés chacun entre deux ailettes (14) directement voisines. Pour créer un segment de canal d'écoulement (12) particulièrement peu encombrant, les éléments de protection (22) sont disposés contre les plates-formes (16) en laissant un interstice intermédiaire et des ouvertures de refroidissement par impact de jets (24) sont ménagées dans la plate-forme (16).
EP10725431A 2009-06-23 2010-06-15 Segment de canal d'écoulement de forme annulaire pour une turbomachine Withdrawn EP2446119A1 (fr)

Priority Applications (1)

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

Applications Claiming Priority (3)

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
PCT/EP2010/058352 WO2010149528A1 (fr) 2009-06-23 2010-06-15 Segment de canal d'écoulement de forme annulaire pour une turbomachine

Publications (1)

Publication Number Publication Date
EP2446119A1 true EP2446119A1 (fr) 2012-05-02

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 Before (1)

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

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)

Families Citing this family (8)

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Publication number Priority date Publication date Assignee Title
US8734111B2 (en) * 2011-06-27 2014-05-27 General Electric Company Platform cooling passages and methods for creating platform cooling passages in turbine rotor blades
EP2634373A1 (fr) 2012-02-28 2013-09-04 Siemens Aktiengesellschaft Agencement pour turbomachine
US11111801B2 (en) * 2013-06-17 2021-09-07 Raytheon Technologies Corporation Turbine vane with platform pad
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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US3433015A (en) * 1965-06-23 1969-03-18 Nasa Gas turbine combustion apparatus
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
US5281097A (en) * 1992-11-20 1994-01-25 General Electric Company Thermal control damper for turbine rotors
JPH08135402A (ja) * 1994-11-11 1996-05-28 Mitsubishi Heavy Ind Ltd ガスタービン静翼構造
FR2758855B1 (fr) * 1997-01-30 1999-02-26 Snecma Systeme de ventilation des plates-formes des aubes mobiles
JP3453293B2 (ja) * 1998-03-03 2003-10-06 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
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
DE50011923D1 (de) * 2000-12-27 2006-01-26 Siemens Ag Gasturbinenschaufel und Gasturbine
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
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Also Published As

Publication number Publication date
US20120100008A1 (en) 2012-04-26
EP2282014A1 (fr) 2011-02-09
JP5443600B2 (ja) 2014-03-19
JP2012530870A (ja) 2012-12-06
CN102803658A (zh) 2012-11-28
WO2010149528A1 (fr) 2010-12-29

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