US20120100008A1 - Annular flow channel section for a turbomachine - Google Patents

Annular flow channel section for a turbomachine Download PDF

Info

Publication number
US20120100008A1
US20120100008A1 US13/379,530 US201013379530A US2012100008A1 US 20120100008 A1 US20120100008 A1 US 20120100008A1 US 201013379530 A US201013379530 A US 201013379530A US 2012100008 A1 US2012100008 A1 US 2012100008A1
Authority
US
United States
Prior art keywords
flow passage
platform
platforms
blade
shielding element
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.)
Abandoned
Application number
US13/379,530
Other languages
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
Assigned to SIEMENS AKTIENGESELLSCHAFT reassignment SIEMENS AKTIENGESELLSCHAFT ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: AHMAD, FATHI
Publication of US20120100008A1 publication Critical patent/US20120100008A1/en
Abandoned legal-status Critical Current

Links

Images

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 refers to an annular flow passage section for a turbomachine, having a stator blade ring which has a number of stator blades arranged in series in the circumferential direction and in each case comprising a blade root, a platform and a blade airfoil which projects radially into the flow passage, the flow passage being delimited on the platform side by shielding elements which are seated in each case between two directly adjacent blade airfoils.
  • annular flow passage section which is referred to in the introduction, is known from EP 1 219 787 B1, for example.
  • the printed patent specification discloses a ring of cast stator blades of an axial flow turbine, in which the stator blades have an aerodynamically curved blade airfoil, provision being made in each case for platforms on the radially outer (root-side) end and radially inner (tip-side) end of the blade airfoil. Installed in the turbine, the platforms are covered by ceramic heat shields.
  • the heat shields are designed in such a way that in pairs they cover a platform half of two directly adjacent stator blades in each case.
  • the heat shields therefore extend basically from the suction-side wall of the blade airfoil of a first stator blade to the pressure-side wall of the blade airfoil of a second stator blade.
  • the ceramic heat shield is fixedly connected in this case, via a tongue, to the gas turbine blade so that the heat shield is fastened with capability of being exchanged.
  • a construction for fastening such a cover, which is alternative to this, is disclosed in US 2007/0237630 A1.
  • Ceramic heat shields require a comparatively large wall thickness in order to be able to durably and reliably withstand the temperatures of the hot gas which occur in a stationary gas turbine. If such ceramic heat shields are used both on the tip-side platform and on the root-side platform of stator blades, this leads to comparatively large turbine stator blades with correspondingly increased space requirement, which also increases the production costs.
  • a modular turbine blade with two sheet-metal covers which, in addition to the associated platform half, also covers the transition to the aerodynamically curved blade airfoil in each case, is known from EP 1 557 535 A1.
  • the sealing of the gap between abutting platform halves of adjacent turbine blades by means of a sealing element inserted in grooves is disadvantageous in this case, however.
  • a form of delimiting the flow passage, which differs from this, follows from EP 1 557 534 A1. It is shown therein that a platform half of the arrangement known from EP 1 557 535 A1 can be dispensed with if one of the sheet-metal covers is supported on the adjacent turbine blade. For that reason, one platform half can be dispensed with. In the case of this development, a close contact of the sheet-metal cover on the adjacent turbine blade is not always ensured however.
  • the object of the invention is therefore the provision of an annular flow passage section for a turbomachine, which requires comparatively little space and, furthermore, guides the hot gas flowing in the flow passage section in a particularly reliable and safe manner, for a particularly long period, without premature wear phenomena occurring on the components which border on the flow passage.
  • the object is achieved with an annular flow passage section for a turbomachine, in which the shielding element is arranged on the platforms, forming a gap, and impingement cooling holes are provided in the platform for impingement cooling of the shielding elements.
  • the invention is based on the knowledge that the platform halves, which are fanned on the stator blades, can also then be protected against the hot gas and its corrosive and thermal influences if the shielding element does not consist of ceramic. In this case, the shielding element is then to be cooled to an adequate degree.
  • impingement cooling of the shielding element is used.
  • this can be designed with thinner walls than in the case of the prior art.
  • the comparatively thin-walled design of the shielding element saves space and is also more cost-effective.
  • the blade airfoil of corresponding stator blades can consequently be of a shorter design in its span without reducing the flow cross section of the annular flow passage section compared with the flow passage section which is known from the prior art.
  • Stator blades which are used in the flow passage section according to the invention, are customarily produced in a casting process and are therefore mainly in one piece. Since the platforms or platform halves of such stator blades previously had to withstand not only the pressure of the hot gas but also had to transmit the mechanical load—caused by the flow forces—of the blade airfoil onto a rear-side hook-fastening, these previously had comparatively solid walls, i.e. large wall thicknesses. This led to poor coolability of platforms, as a result of which the service life of such stator blades was previously also limited by the platforms. By using a shielding element according to the invention, especially the thermal load of such platforms can be reduced, which leads to a significant extension of the service life of stator blades.
  • the transition can now be better protected against the direct contact and influence of the hot gas which flows in the flow passage since at this point there is now a gap between the shielding element and the blade airfoil wall or transition, through which gap the cooling medium which is used for the impingement cooling of the shielding element—for example cooling air after completion of the impingement cooling—can discharge into the flow passage. Also, this leads to an extended service life of the stator blade due to the reduction of the thermal load in the region of the transition from the platform to the blade airfoil.
  • each shielding element extends across a gap which is delimited by the platforms of two directly adjacent stator blades. This enables a low-loss guiding of the hot gas in the flow passage, even in the event that a displacement of mutually adjacent platforms occurs on account of thermally induced expansions.
  • the shielding elements preferably have in each case a baseplate, consisting of a metallic material, which delimits the flow passage and is produced separately from the stator blades.
  • a baseplate consisting of a metallic material, which delimits the flow passage and is produced separately from the stator blades.
  • metallic materials As a result of the cooling of the shielding element, recourse can be made to metallic materials.
  • the entire shielding element is produced separately from the stator blades. This has the advantage that in the event of wear phenomena occurring on the shielding element only this is to be replaced and not the complete stator blade, as in the case of unshielded stator blade platforms.
  • the shielding element is preferably produced from a metallic material with good insulating properties.
  • the wall thickness of the baseplate is less than the wall thickness of the platform which is covered by the shielding element.
  • a flow passage section which is compact in respect to space requirement can be disclosed, which reduces the production costs and material costs for such a flow passage section.
  • the shielding element has a protective coating, especially a thermal barrier coating, on the flow passage side.
  • FIG. 1 shows a section through two of the blade airfoils of an annular flow passage section as a developed view thereof, with a shielding element arranged over the platforms of the stator blades, and
  • FIG. 2 shows the section according to the line of intersection II-II through the platform of the stator blade and through the shielding element.
  • FIG. 1 shows the cross section through the blade airfoils 14 of two stator blades 10 of an annular flow passage section 12 of a turbomachine, for example a gas turbine, through which hot gas can axially flow.
  • the flow passage section 12 essentially comprises a stator blade ring with a large number of stator blades 10 which are arranged in series in the circumferential direction.
  • the stator blades 10 in this case are fastened on a stator blade carrier in a conventional manner. The view in FIG.
  • a shielding element 22 is arranged in a form-fitting manner between a suction-side blade airfoil wall 18 of the stator blade 10 which is shown further down in FIG. 1 and the pressure-side blade airfoil wall 20 of the stator blade 10 which is shown further up in FIG. 1 .
  • the shielding element 22 is essentially, i.e. on the hot gas side, formed in one piece and completely covers the subjacent halves of the platforms 16 between the blade airfoils 14 of the two directly adjacent stator blades 10 .
  • the stator blade ring has such a shielding element 22 between each pair of directly adjacent blade airfoils 14 in each case, wherein adjacent shielding elements 22 , moreover, on one side butt against each other upstream of a leading edge 21 of the blade airfoil 14 and downstream of a trailing edge 23 of the blade airfoil 14 with a gap which is as small as possible.
  • impingement cooling holes 24 are arranged in the platforms 16 , for example in a grid-like manner.
  • FIG. 2 shows the section according to the line of intersection II-II through the stator blade 10 and the shielding element 22 .
  • the shielding element 22 is arranged on the platform 16 on the hot gas side, forming a gap, provision being made in the platform 16 for impingement cooling holes 24 which extend obliquely, for example, to its surface.
  • a cooling medium K is fed to the rear space 28 facing away from the flow passage 26 , which cooling medium can discharge from the rear space 28 through the impingement cooling holes 24 and in a jet-like manner can enter the gap between the shielding element 22 and the platform 16 . With the impingement of the impingement cooling jets, these cool the shielding element 22 so that despite the hot gas which flows through the flow passage 26 this shielding element has an adequate service life.
  • the shielding element 22 which is shown in cross section in FIG. 2 consists of metal and essentially comprises a baseplate 30 which extends parallel to the passage-side platform surface. On the two edges of the baseplate 30 which lie opposite each other, provision is made for wall sections 32 which at the sides project transversely to the baseplate 30 and encompass corresponding sidewalls of the platform 16 in a clamp-like manner.
  • the wall thickness of the baseplate 30 is significantly less in this case than the wall thickness of the platform 16 in the region of the impingement cooling holes 24 .
  • the shielding element 22 on the stator blade 10 or on the platform 16 this can be screwed on, for example, as shown by the dash-dot lines.
  • Other types of fastening such as clamping, especially form-fitting clamping of the shielding element 22 on the platform 16 , are also conceivable.
  • the shielding element 22 on its surface which is exposed to the hot gas, can have a thermal barrier coating in order to further increase its thermal resistance.
  • the cooling medium K which flows into the gap between the shielding element 22 and the platform surface, flows out at that gap 36 ( FIG. 1 ) which is provided between the shielding element 22 and the suction-side blade airfoil wall 18 or pressure-side blade airfoil wall 20 after impingement cooling has been carried out.
  • the platform 16 which is shown in FIG. 2 and the shielding element 22 which is arranged above it in this case can be both a root-side platform and a tip-side platform of stator blades 10 , providing the stator blades 10 which are used in the annular flow passage section 12 have platforms 16 which extend transversely to the blade airfoil 14 at both opposite ends of the blade airfoil 14 .
  • the invention can also be used on only one of the two platforms 16 of such a stator blade 10 .
  • annular flow passage section 12 for a turbomachine having a stator blade ring which has a number of stator blades 10 arranged in series in the circumferential direction and in each case comprising a platform 16 , and a blade airfoil 14 which projects radially into the flow passage 26 , wherein the flow passage 26 , on the platform side, is delimited by shielding elements 22 which are arranged in each case between two directly adjacent blade airfoils 14 , wherein for forming a particularly space-saving flow passage section 12 the shielding elements 22 are arranged on the platforms 16 , forming a gap, and impingement cooling holes 24 are provided in the platform 16 .

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US13/379,530 2009-06-23 2010-06-15 Annular flow channel section for a turbomachine Abandoned US20120100008A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP09008227A EP2282014A1 (de) 2009-06-23 2009-06-23 Rinförmiger Strömungskanalabschnitt für eine Turbomaschine
EP09008227.2 2009-06-23
PCT/EP2010/058352 WO2010149528A1 (de) 2009-06-23 2010-06-15 Ringförmiger strömungskanalabschnitt für eine turbomaschine

Publications (1)

Publication Number Publication Date
US20120100008A1 true US20120100008A1 (en) 2012-04-26

Family

ID=41351921

Family Applications (1)

Application Number Title Priority Date Filing Date
US13/379,530 Abandoned US20120100008A1 (en) 2009-06-23 2010-06-15 Annular flow channel section for a turbomachine

Country Status (5)

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

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9863271B2 (en) 2012-02-28 2018-01-09 Siemens Aktiengesellschaft Arrangement for a turbomachine
EP3323992A1 (en) * 2016-11-17 2018-05-23 United Technologies Corporation Airfoil with panel and side edge cooling
EP3663540A1 (en) * 2018-12-07 2020-06-10 United Technologies Corporation Cooling system
CN112943378A (zh) * 2021-02-04 2021-06-11 大连理工大学 一种涡轮叶片枝网式冷却结构

Families Citing this family (4)

* Cited by examiner, † Cited by third party
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
JP6247385B2 (ja) * 2013-06-17 2017-12-13 ユナイテッド テクノロジーズ コーポレイションＵｎｉｔｅｄ Ｔｅｃｈｎｏｌｏｇｉｅｓ Ｃｏｒｐｏｒａｔｉｏｎ プラットフォームパッドを備えるタービンベーン
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

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3628880A (en) * 1969-12-01 1971-12-21 Gen Electric Vane assembly and temperature control arrangement
US5281097A (en) * 1992-11-20 1994-01-25 General Electric Company Thermal control damper for turbine rotors
US6017189A (en) * 1997-01-30 2000-01-25 Societe National D'etede Et De Construction De Moteurs D'aviation (S.N.E.C.M.A.) Cooling system for turbine blade platforms
US6457935B1 (en) * 2000-06-15 2002-10-01 Snecma Moteurs System for ventilating a pair of juxtaposed vane platforms
US20050196278A1 (en) * 2004-03-06 2005-09-08 Rolls-Royce Plc Turbine blade arrangement
US8240987B2 (en) * 2008-08-15 2012-08-14 United Technologies Corp. Gas turbine engine systems involving baffle assemblies

Family Cites Families (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3433015A (en) * 1965-06-23 1969-03-18 Nasa Gas turbine combustion apparatus
US4218178A (en) * 1978-03-31 1980-08-19 General Motors Corporation Turbine vane structure
JPH08135402A (ja) * 1994-11-11 1996-05-28 Mitsubishi Heavy Ind Ltd ガスタービン静翼構造
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
WO1999054597A1 (de) * 1998-04-21 1999-10-28 Siemens Aktiengesellschaft Turbinenschaufel
WO1999060253A1 (de) * 1998-05-18 1999-11-25 Siemens Aktiengesellschaft Gekühlte turbinenschaufelplattform
DE50011923D1 (de) * 2000-12-27 2006-01-26 Siemens Ag Gasturbinenschaufel und Gasturbine
EP1557534A1 (de) * 2004-01-20 2005-07-27 Siemens Aktiengesellschaft Turbinenschaufel und Gasturbine mit einer solchen Turbinenschaufel
EP1557535A1 (de) * 2004-01-20 2005-07-27 Siemens Aktiengesellschaft Turbinenschaufel und Gasturbine mit einer solchen Turbinenschaufel
DE502005010381D1 (de) * 2005-04-28 2010-11-25 Siemens Ag Verfahren und Vorrichtung zur Einstellung eines Radialspaltes eines axial durchströmten Verdichters einer Strömungsmaschine
US7604456B2 (en) * 2006-04-11 2009-10-20 Siemens Energy, Inc. Vane shroud through-flow platform cover
US7488157B2 (en) * 2006-07-27 2009-02-10 Siemens Energy, Inc. Turbine vane with removable platform inserts
US7766609B1 (en) * 2007-05-24 2010-08-03 Florida Turbine Technologies, Inc. Turbine vane endwall with float wall heat shield

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3628880A (en) * 1969-12-01 1971-12-21 Gen Electric Vane assembly and temperature control arrangement
US5281097A (en) * 1992-11-20 1994-01-25 General Electric Company Thermal control damper for turbine rotors
US6017189A (en) * 1997-01-30 2000-01-25 Societe National D'etede Et De Construction De Moteurs D'aviation (S.N.E.C.M.A.) Cooling system for turbine blade platforms
US6457935B1 (en) * 2000-06-15 2002-10-01 Snecma Moteurs System for ventilating a pair of juxtaposed vane platforms
US20050196278A1 (en) * 2004-03-06 2005-09-08 Rolls-Royce Plc Turbine blade arrangement
US8240987B2 (en) * 2008-08-15 2012-08-14 United Technologies Corp. Gas turbine engine systems involving baffle assemblies

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9863271B2 (en) 2012-02-28 2018-01-09 Siemens Aktiengesellschaft Arrangement for a turbomachine
EP3323992A1 (en) * 2016-11-17 2018-05-23 United Technologies Corporation Airfoil with panel and side edge cooling
US10598029B2 (en) 2016-11-17 2020-03-24 United Technologies Corporation Airfoil with panel and side edge cooling
US11319817B2 (en) 2016-11-17 2022-05-03 Raytheon Technologies Corporation Airfoil with panel and side edge cooling
EP3663540A1 (en) * 2018-12-07 2020-06-10 United Technologies Corporation Cooling system
CN112943378A (zh) * 2021-02-04 2021-06-11 大连理工大学 一种涡轮叶片枝网式冷却结构

Also Published As

Publication number Publication date
JP5443600B2 (ja) 2014-03-19
EP2446119A1 (de) 2012-05-02
EP2282014A1 (de) 2011-02-09
CN102803658A (zh) 2012-11-28
JP2012530870A (ja) 2012-12-06
WO2010149528A1 (de) 2010-12-29

Similar Documents

Publication Publication Date Title
US20120100008A1 (en) Annular flow channel section for a turbomachine
US7044710B2 (en) Gas turbine arrangement
US8186965B2 (en) Recovery tip turbine blade
EP1749975B1 (en) Cooled turbine shroud
EP1759089B1 (en) Gas turbine blade shroud
US6059530A (en) Twin rib turbine blade
US10408073B2 (en) Cooled CMC wall contouring
US9175579B2 (en) Low-ductility turbine shroud
US7766609B1 (en) Turbine vane endwall with float wall heat shield
US6273682B1 (en) Turbine blade with preferentially-cooled trailing edge pressure wall
US7988410B1 (en) Blade tip shroud with circular grooves
EP1748155B1 (en) Cooled shroud assembly and method of cooling a shroud
EP2586996B1 (en) Turbine bucket angel wing features for forward cavity flow control and related method
EP2631434A2 (en) Low-ductility turbine shroud
RU2494263C2 (ru) Лопатки лопаточного колеса газотурбинного двигателя, оснащенные канавками для охлаждения
US8118547B1 (en) Turbine inter-stage gap cooling arrangement
EP2597264B1 (en) Aerofoil cooling arrangement
US20170183971A1 (en) Tip shrouded turbine rotor blades
US8061989B1 (en) Turbine blade with near wall cooling
US8382424B1 (en) Turbine vane mate face seal pin with impingement cooling
US9745852B2 (en) Axial rotor portion and turbine rotor blade for a gas turbine
US20120201694A1 (en) Turbine blade
US20100068069A1 (en) Turbine Blade
US20110038708A1 (en) Turbine endwall cooling arrangement
EP2169183B1 (en) Turbine nozzle with curved recesses in the outer platforms

Legal Events

Date Code Title Description
AS Assignment

Owner name: SIEMENS AKTIENGESELLSCHAFT, GERMANY

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:AHMAD, FATHI;REEL/FRAME:027420/0026

Effective date: 20111130

STCB Information on status: application discontinuation

Free format text: ABANDONED -- FAILURE TO PAY ISSUE FEE