US20130266427A1 - Sealing system for a turbomachine - Google Patents

Sealing system for a turbomachine Download PDF

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Publication number
US20130266427A1
US20130266427A1 US13/856,535 US201313856535A US2013266427A1 US 20130266427 A1 US20130266427 A1 US 20130266427A1 US 201313856535 A US201313856535 A US 201313856535A US 2013266427 A1 US2013266427 A1 US 2013266427A1
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United States
Prior art keywords
sealing
disposed
sealing system
flow
accordance
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
US13/856,535
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English (en)
Inventor
Inga Mahle
Alexander Boeck
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.)
MTU Aero Engines AG
Original Assignee
MTU Aero Engines GmbH
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Filing date
Publication date
Application filed by MTU Aero Engines GmbH filed Critical MTU Aero Engines GmbH
Assigned to MTU AERO ENGINES GMBH reassignment MTU AERO ENGINES GMBH ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: BOECK, ALEXANDER, MAHLE, INGA
Publication of US20130266427A1 publication Critical patent/US20130266427A1/en
Abandoned legal-status Critical Current

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    • 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
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/122—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable material
    • 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
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • 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
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/127—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with a deformable or crushable structure, e.g. honeycomb
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • 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
    • F05D2240/55—Seals
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the invention relates to a sealing system for a turbomachine, in particular, for a gas turbine, wherein the sealing system is disposed in an annular space between a flow-restricting wall of the turbomachine and at least one series of rotor blades comprising a plurality of rotor blades, and at least two seals.
  • the invention furthermore relates to a gas turbine, in particular, an aircraft engine that includes at least one sealing system.
  • Sealing systems of this type are employed, in particular, in connection with so-called clearance-maintaining systems in the components of compressors and turbines. These sealing systems function to keep to a minimum a sealing gap of the rotating blading relative to a housing, as well as the clearances of a stationary blading relative to a rotating rotor hub, thereby ensuring stable operating characteristics together with high efficiency.
  • the rotating components of the turbine typically include sealing fins that in the known way touch or run in against honeycomb seals.
  • the seals are provided here in the form of rub linings and abradable linings Sealing systems of this type are disclosed, for example, in U.S. Pat. No. 4,856,963 B1 and DE 198 07 247 A1.
  • the object of this invention is therefore to create a sealing system of the type referenced above that reliably provides an increase in the efficiency of a turbomachine.
  • a further object of the invention is to provide a corresponding gas turbine that exhibits improved efficiency.
  • a sealing system according to the invention for a turbomachine is disposed in an annular space between a flow-restricting wall of the turbomachine and at least one series of rotor blades comprising a plurality of rotor blades.
  • the sealing system here comprises at least two seals, wherein the first seal comprises at least one first sealing fin that is disposed on an end of the rotor blade facing the wall as well as at least one first abradable lining that is disposed on an inside of the wall and that is opposite the first sealing fin, and the second seal is disposed in the direction of flow following the first seal, and comprises a second abradable lining that is provided on the end of the rotor blade facing the wall as well as a second sealing fin that is disposed on the inside of the wall and is opposite the second abradable lining.
  • the first sealing fin and the second abradable lining here can be disposed on a shroud of the rotor blade.
  • this can be a casing of the turbomachine.
  • the sealing system according to the invention is able to ensure that any leakage flow is conducted closely along the end of the rotor blade facing the wall or along the shroud of the rotor blade until it reenters the so-called main flow of the turbomachine, with the result that this is delayed less severely until it reenters the main flow as compared with conventional sealing systems, in particular, sealing systems that have at least two sealing fins on the shroud of the rotor blade.
  • the leakage flow can be introduced with less disruption, that is, with less generation of turbulence, into the main flow.
  • the seal of the sealing system according to the invention which seal is disposed last in the flow direction, is always of the same design as the referenced second seal in order to achieve the described advantages. Otherwise the sealing system according to the invention can be implemented using already known techniques.
  • the sealing system can thus be of relatively lightweight construction so as to prevent any failure of the bond between the second abradable lining and the rotor blade from causing the entire sealing system to fail since those components of the sealing system that are located further forward in the flow direction, that is, the first seal, continue to be present. Also minimized is any secondary damage, such as, for example, that caused by the components of the second abradable lining impacting the stages of the turbomachine located further back in the flow direction, due to the relatively small size of the second abradable lining.
  • the second abradable lining has at least one chamfer in a rearward region in the flow direction.
  • the shroud can have at least one chamfer in a rearward region in the flow direction.
  • the chamfer here can be at an angle a ranging between 10° and 65°, preferably, approximately 20° and 45°, relative to a longitudinal axis of the turbomachine.
  • the first and/or second sealing fin is tilted opposite the flow direction. This also enables turbulences to be significantly reduced within the leakage flow.
  • the first sealing fin is integrated into the shroud. It is also possible to integrate the second sealing fin in the flow-restricting wall. This yields production advantages that result in overall reduced manufacturing costs.
  • the sealing system is provided in the form of a stepped labyrinth. This allows the sealing system to be advantageously adapted to the relevant specific requirements of the turbomachine, in particular, to the requirements presented by a low-pressure turbine.
  • the invention furthermore relates to a gas turbine, in particular, an aircraft engine, comprising at least one sealing system, wherein the sealing system is disposed in an annular space between a flow-restricting wall of the gas turbine and at least one series of rotor blades comprising a plurality of rotor blades, and at least two seals, wherein the first seal comprises at least one sealing fin that is disposed on an end of the rotor blade facing the wall as well as at least one first abradable lining that is disposed on an inside of the wall and is opposite the first sealing fin, and the second seal is disposed in the flow direction following the first seal and comprises a second abradable lining that is provided on the end of the rotor blade facing the wall as well as a second sealing fin that is disposed on the inside of the wall and is opposite the second abradable lining.
  • the first sealing fin and the second abradable lining here can be disposed on a shroud of the rotor blade.
  • the flow-restricting wall can be a casing of the turbomachine.
  • the gas turbine according to the invention exhibits improved efficiency due to the fact that a leakage flow can be conducted by the sealing system closely along the end of the rotor blades facing the flow-restricting wall up to re-entry into a main flow of the gas turbine such that this maintains a high velocity up to the point of re-entry, with the result that the velocity difference vis-à-vis the relatively fast main flow is reduced.
  • the referenced sealing system ensures low generation of turbulence in the leakage flow, thereby allowing this flow to be introduced into the main flow in relatively trouble-free fashion. In overall terms, this results in a reduction in the referenced mixing losses and in a more advantageous flow around the following blade.
  • the seal that is disposed last in the flow direction is always of the same design as the referenced second seal.
  • the sealing system for the gas turbine according to the invention can otherwise be implemented using already known techniques.
  • the sealing system can thus be of relatively lightweight construction so that any failure of the bond between the second abradable lining and the rotor blade does not result in failure of the entire sealing system due to the fact that the components of the sealing system located further forward in the flow direction continue to be present.
  • any consequential damage such as, for example, that caused by components of the second abradable lining impacting stages of the turbomachine located further back in the flow direction are minimized due to the relatively small size of the second abradable lining. Also minimized is the risk of a rotor imbalance in the turbomachine caused by asymmetric rubbing in the gas turbine according to the invention, or by the sealing system according to the invention.
  • FIGURE here provides a sectional, schematic, and partial cutaway view of a sealing system according to the invention.
  • the sealing system 10 depicted in the figure is a component of a low-pressure gas turbine. It is evident that sealing system 10 is disposed in an annular space 30 between a flow-restricting wall 22 of the low-pressure gas turbine and a series of rotor blades comprising a plurality of rotor blades 12 . Sealing system 10 here comprises two seals 34 , 36 , second seal 36 being disposed following first seal 34 in the flow direction 28 .
  • First seal 34 comprises a first sealing fin 18 that is provided on a shroud 16 of rotor blade 12 , and a first abradable lining 24 that is disposed on an inside of wall 22 and is opposite first sealing fin 18 .
  • Second seal 36 comprises a second abradable lining 20 that is disposed following first sealing fin 18 in flow direction 28 , and a second sealing fin 26 that is disposed on the inside of wall 22 , which sealing fin is disposed following first abradable lining 24 in flow direction 28 and is opposite second abradable lining 20 .
  • First and second rub and abradable linings 24 , 20 can be provided in the conventional form. They have a honeycomb structure in the embodiment shown. Another possible approach is to produce first and/or second abradable linings 24 , 20 by spraying them onto shroud 16 or the inside of wall 22 . It is possible, in particular, to employ powder injection molding for this purpose.
  • first and/or second abradable linings 24 , 20 are disposed as a complete system in the referenced regions or to bond them to these regions, for example, by brazing.
  • Second abradable lining 20 in the illustrated embodiment is disposed axially following a reinforcement rib 38 that is created by a Z-interlocking of shrouds 16 . This advantageously creates a flat attachment surface for second abradable lining 20 on the surface of shroud 16 .
  • second abradable lining 20 has at least one chamfer 32 in rearward region in flow direction 28 .
  • Chamfer 32 in the illustrated embodiment here is at an angle a less than 30° relative to a longitudinal axis of the turbomachine, for example, of the low-pressure turbine. It is possible in general for angle to lie within a range between 10° and 65°, preferably, a range between 20° and 45°.
  • sealing system 10 is provided in the form of a stepped labyrinth.
  • First and second abradable linings 24 , 20 , as well as first and second sealing fins 18 , 26 of sealing system 10 can be composed of such materials as are typically used. A wide variety of materials is known that can be used here.
  • First sealing fin 18 in the illustrated embodiment is integrated into shroud 16 .
  • Second sealing fin 26 is attached to the inside of wall 22 .
  • a guide vane 14 is disposed following rotor blade 12 in the flow direction.
  • the sealing system described in the embodiment is not restricted to the area of low-pressure turbines.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Sealing Using Fluids, Sealing Without Contact, And Removal Of Oil (AREA)
US13/856,535 2012-04-04 2013-04-04 Sealing system for a turbomachine Abandoned US20130266427A1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EP12163065.1A EP2647796A1 (fr) 2012-04-04 2012-04-04 Système d'étanchéité pour turbomachine
EP12163065.1 2012-04-04

Publications (1)

Publication Number Publication Date
US20130266427A1 true US20130266427A1 (en) 2013-10-10

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Family Applications (1)

Application Number Title Priority Date Filing Date
US13/856,535 Abandoned US20130266427A1 (en) 2012-04-04 2013-04-04 Sealing system for a turbomachine

Country Status (2)

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US (1) US20130266427A1 (fr)
EP (1) EP2647796A1 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102016102923A1 (de) * 2016-02-19 2017-08-24 Abb Turbo Systems Ag Verdichterradrückseitige Radialturbine
US10415735B2 (en) 2015-06-17 2019-09-17 Rolls-Royce Corporation Labyrinth seal with tunable flow splitter
CN114207252A (zh) * 2019-08-06 2022-03-18 赛峰飞机发动机公司 用于涡轮机涡轮并包括设有导向叶片的磨损面的可磨损部件
US11339676B2 (en) 2017-12-28 2022-05-24 Mitsubishi Heavy Industries Aero Engines, Ltd. Aircraft gas turbine, and rotor blade of aircraft gas turbine
US20220259983A1 (en) * 2020-03-30 2022-08-18 Ihi Corporation Secondary flow suppression structure

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9598981B2 (en) * 2013-11-22 2017-03-21 Siemens Energy, Inc. Industrial gas turbine exhaust system diffuser inlet lip
DE102016222720A1 (de) * 2016-11-18 2018-05-24 MTU Aero Engines AG Dichtungssystem für eine axiale Strömungsmaschine und axiale Strömungsmaschine

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6139263A (en) * 1998-02-20 2000-10-31 Klingels; Hermann Flow machine with rotor and stator
US7070387B2 (en) * 2001-08-30 2006-07-04 Snecma Moteurs Gas turbine stator housing
US20080056895A1 (en) * 2006-08-31 2008-03-06 Shigeki Senoo Axial turbine
WO2011054341A2 (fr) * 2009-11-07 2011-05-12 Mtu Aero Engines Gmbh Ensemble d'étanchéité pour une turbine à gaz et turbine à gaz de ce type
US20120230818A1 (en) * 2009-09-30 2012-09-13 Andrew Shepherd Airfoil and corresponding guide vane, blade, gas turbine and turbomachine
US20130272888A1 (en) * 2012-04-13 2013-10-17 General Electric Company Turbomachine blade tip shroud with parallel casing configuration

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US4856963A (en) 1988-03-23 1989-08-15 United Technologies Corporation Stator assembly for an axial flow rotary machine
EP1152124A1 (fr) * 2000-05-04 2001-11-07 Siemens Aktiengesellschaft Garniture d'étanchéité
US7255531B2 (en) * 2003-12-17 2007-08-14 Watson Cogeneration Company Gas turbine tip shroud rails
JP4668976B2 (ja) * 2007-12-04 2011-04-13 株式会社日立製作所 蒸気タービンのシール構造
US7909335B2 (en) * 2008-02-04 2011-03-22 General Electric Company Retractable compliant plate seals
DE102009040758A1 (de) * 2009-09-10 2011-03-17 Mtu Aero Engines Gmbh Umlenkvorrichtung für einen Leckagestrom in einer Gasturbine und Gasturbine
US20110070072A1 (en) * 2009-09-23 2011-03-24 General Electric Company Rotary machine tip clearance control mechanism

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6139263A (en) * 1998-02-20 2000-10-31 Klingels; Hermann Flow machine with rotor and stator
US7070387B2 (en) * 2001-08-30 2006-07-04 Snecma Moteurs Gas turbine stator housing
US20080056895A1 (en) * 2006-08-31 2008-03-06 Shigeki Senoo Axial turbine
US20120230818A1 (en) * 2009-09-30 2012-09-13 Andrew Shepherd Airfoil and corresponding guide vane, blade, gas turbine and turbomachine
WO2011054341A2 (fr) * 2009-11-07 2011-05-12 Mtu Aero Engines Gmbh Ensemble d'étanchéité pour une turbine à gaz et turbine à gaz de ce type
US20130272888A1 (en) * 2012-04-13 2013-10-17 General Electric Company Turbomachine blade tip shroud with parallel casing configuration

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
English Machine Translation of WO2011054341A2 which was translated from the EPO on 11/19/2016. *

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10415735B2 (en) 2015-06-17 2019-09-17 Rolls-Royce Corporation Labyrinth seal with tunable flow splitter
DE102016102923A1 (de) * 2016-02-19 2017-08-24 Abb Turbo Systems Ag Verdichterradrückseitige Radialturbine
US11339676B2 (en) 2017-12-28 2022-05-24 Mitsubishi Heavy Industries Aero Engines, Ltd. Aircraft gas turbine, and rotor blade of aircraft gas turbine
CN114207252A (zh) * 2019-08-06 2022-03-18 赛峰飞机发动机公司 用于涡轮机涡轮并包括设有导向叶片的磨损面的可磨损部件
US20220259983A1 (en) * 2020-03-30 2022-08-18 Ihi Corporation Secondary flow suppression structure
US11808156B2 (en) * 2020-03-30 2023-11-07 Ihi Corporation Secondary flow suppression structure

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Publication number Publication date
EP2647796A1 (fr) 2013-10-09

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