WO2011054341A2 - Ensemble d'étanchéité pour une turbine à gaz et turbine à gaz de ce type - Google Patents

Ensemble d'étanchéité pour une turbine à gaz et turbine à gaz de ce type Download PDF

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Publication number
WO2011054341A2
WO2011054341A2 PCT/DE2010/001279 DE2010001279W WO2011054341A2 WO 2011054341 A2 WO2011054341 A2 WO 2011054341A2 DE 2010001279 W DE2010001279 W DE 2010001279W WO 2011054341 A2 WO2011054341 A2 WO 2011054341A2
Authority
WO
WIPO (PCT)
Prior art keywords
sealing
sealing element
flow
gas turbine
leakage
Prior art date
Application number
PCT/DE2010/001279
Other languages
German (de)
English (en)
Other versions
WO2011054341A3 (fr
Inventor
Inga Mahle
Original Assignee
Mtu Aero Engines Gmbh
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 Mtu Aero Engines Gmbh filed Critical Mtu Aero Engines Gmbh
Publication of WO2011054341A2 publication Critical patent/WO2011054341A2/fr
Publication of WO2011054341A3 publication Critical patent/WO2011054341A3/fr

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
    • F01D11/08—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
    • F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • 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
    • F05D2240/10—Stators
    • F05D2240/11—Shroud seal segments
    • 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

Definitions

  • the invention relates to a sealing arrangement for a gas turbine, in particular for an aircraft gas turbine, according to the preamble of patent claim 1 and a gas turbine with such a sealing arrangement according to the preamble of patent claim 12.
  • the flow direction of the main flow passes through an array of guide and blade rows.
  • the outer engine shell forms a housing in which the guide blade rows are radially centered and mounted axially fixed.
  • the outer sealing arrangement is designed as a labyrinth seal, wherein annular cutting-like sealing tips as sealing elements run against honeycomb structures (honeycombs), which are applied to honeycomb carriers.
  • honeycomb carriers are in turn mounted on the guide blade rows and connected via this with the housing.
  • the sealing elements are arranged in an annular space (outer cavity) between the housing and in each case one blade row, wherein the annular space is traversed by a leakage flow, which then opens again into the main flow.
  • the object of the invention is to provide a sealing arrangement and a gas turbine with such a sealing arrangement in which the aforementioned disadvantages are eliminated and a reduction of the mixing losses is achieved.
  • a sealing arrangement according to the invention for a gas turbine in particular for an aircraft gas turbine, has at least one sealing element, wherein the sealing element is arranged in an annular space between a housing and at least one row of blades, in particular a blade row, and wherein the annular space is traversed by a leakage flow, which flows from a main stream is branched off and after the annulus at a reentry point opens into the main stream.
  • the at least one sealing element is arranged in a particularly preferred embodiment in the region of a blade row, wherein the leakage flow flows in front of the blades following stator blades in the main flow.
  • At least one sealing element is assigned to the housing, ie fixed directly or indirectly to the housing and extends in the direction of the blade row, wherein a leakage point between the sealing element and the blade row is formed in the region of the reentry point of the leakage flow into the main flow.
  • the leakage point for example the leakage gap in a contact-free sealing element or the leakage surface in a brush seal in relation to the prior art according to DE 198 07 247 A1
  • the leakage flow remains until reentry in the main stream at a high speed, so that the speed gradients are reduced to relatively fast main flow.
  • the leakage flow is conducted closely along the shroud until reentry into the main stream and may be low in interference, i. with less vortex binding, are introduced into the main stream. Overall, this leads to a reduction of the mixing losses and an improved oncoming flow of the subsequent blade row, in particular in the wall area.
  • At least one sealing element is arranged in a rear region in the direction of flow relative to the blade row.
  • the leakage point is close to the reentry point of the leakage flow into the main flow, so that the leakage flow until reentry into the main flow has a high speed and the velocity gradients are reduced to relatively fast main flow.
  • the blade row can be provided on the housing side with a shroud and / or a squealer or inlet lining, whereby the shroud and / or the squealer or enamel coating in the direction of the main flow starting from a sealing element, preferably a downstream sealing element, is reduced.
  • the shroud and / or the squealer or enema coating is preferably provided with at least one bevel in a rear region in the flow direction, which has a sloping surface sloping down in the flow direction.
  • the leakage flow is conducted closely to the shroud back to the reentry into the main stream and can be introduced by the favorable shape with increasing in the flow direction cross section of the annulus, even less disturbance with further reduced vortex formation in the main stream.
  • the chamfer extends in a preferred exemplary embodiment steadily in the direction of the main flow, so that an inclined surface is formed. It is advantageous if the bevel has an angle in the range of about 10 to 65 °, preferably from about 20 to 45 °, to a longitudinal axis of the gas turbine.
  • the chamfering preferably begins in the region of the tip of a sealing element, preferably in the region of the tip of a downstream sealing element, so that an improved, homogeneous transition of the leakage flow into the main flow is made possible.
  • sealing arrangement at least two axially spaced-apart sealing elements of a row of blades extending essentially in the circumferential direction and designed as sealing tips (sealing membranes) are assigned.
  • sealing tips sealing membranes
  • the shroud and / or the squint or inlet lining may be formed in steps, wherein the sealing elements are preferably assigned to one stage.
  • a gas turbine according to the invention uses at least one such sealing arrangement, which has at least one sealing element, wherein the Sealing element is disposed in an annular space between a housing and at least one row of blades, in particular a blade row, and wherein the annular space is flowed through by a leakage flow, which is branched off from a main flow and flows back to the annular space at a reentry point in the main flow.
  • at least one sealing element is associated with the housing and extends in the direction of the blade row, wherein a leakage point between the sealing element and the blade row is formed in the region of the reentry point of the leakage flow into the main flow.
  • FIGURE shows a schematic diagram of a sealing arrangement according to the invention in the region of a low-pressure turbine of an aircraft gas turbine.
  • the FIGURE shows an aircraft gas turbine 1 in the region of an arrangement of guide and blade rows 2a, 2b, 4 of a low-pressure turbine with a sealing arrangement 6 according to the invention.
  • the sealing elements 8a, 8b are arranged in an annular space 12 or an outer cavity between a housing 14 and in each case one row of blades 4.
  • the annular space 12 is traversed by a leakage flow 16, which is branched off from a main stream 18 and after the annulus 12 at a reentry point 20 before, on the blades 4 following, guide or stator blades 2b, flows back into the main flow 18.
  • the flow direction of the main flow 18 and the leakage flow 16 extends in the figure substantially from left to right through the arrangement of guide and blade rows 2a, 2b, 4.
  • the sealing arrangement 6 has two sealing elements 8a, 8b, which are arranged axially spaced from each other and extend substantially in the circumferential direction. Alternatively, at least one sealing element extends obliquely to the longitudinal axis of the Fluggastrubine.
  • two sealing elements 8a, 8b are assigned to a row of blades 4.
  • the sealing elements 8a, 8b are fixed to the housing 14 and extend radially in Direction of the blade row 4, wherein formed as a leakage gap 22 leakage between the rear sealing element 8b and the blade row 4 in the region of the reentry point 20 of the leakage flow 16 is formed in the main stream 18.
  • the front sealing element 8a forms a front leakage gap 24 with the blade row 4.
  • the leakage flow 16 maintains a high velocity until reentry so that the velocity gradients to the relatively fast main flow 18 are reduced. Due to the position of the rear leakage point 22, the leakage flow 16 is passed close to the honeycomb structure 10 until reentry into the main flow 18 and can be introduced into the main flow 18 with little disturbance, as indicated by the streamline 16. Overall, this leads to a reduction of the mixing losses and an improved oncoming flow of the subsequent blade row, in particular in the wall area.
  • the cross section of the annular space 12 is increased in the direction of the main flow, starting from the rear sealing element 8b.
  • the blade row 4 is the housing side provided with a shroud 26 on which the
  • Honeycomb structure 10 is provided as a squint or inlet lining, which is designed such that the annular space 12 increases in the direction of the main flow 15, starting from the rear in the flow direction sealing element 8b. In other words - the squeaking or inlet lining 10 is reduced in the direction of the main flow 18, starting from the downstream sealing element 8b.
  • the squishing or enamel coating 10 is step-shaped and has peripheral stepped surfaces 28a, 28b.
  • the inlet-side sealing element 8a is assigned to the inlet-side step surface 28a and the outlet-side sealing element 8b to the outlet-side step surface 28b.
  • the step surfaces 28a, 28b extend parallel to the longitudinal axis of the aircraft gas turbine 1, wherein the outlet-side step surface 28b opposite the inlet-side step surface 28a is arranged radially outboard.
  • the squish or enamel coating 10 may be formed continuously, so that the two step surfaces 28a, 28b extend at a common height.
  • the squishing or enema coating 10 is provided in the illustrated embodiment in a rear region in the flow direction with a chamfer 30 having a in Has flow direction sloping inclined surface 32.
  • the inclined surface 32 adjoins the exit-side step surface 28b, so that the leakage flow 16 is passed close to the rubbing or inlet lining 10 until it re-enters the main flow 18. Due to this favorable shaping, the leakage flow 16 can still be introduced into the main flow 18 with less interference with further reduced vortex formation.
  • the chamfer 30 begins in the area of the blade-side tip of the rear sealing element 8b, so that an improved, homogeneous transition of the leakage flow 16 into the main flow 18 is achieved. It is advantageous if the bevel 30 has an angle ⁇ in the range of about 10 to 65 ° to the longitudinal axis of the gas turbine. In the illustrated embodiment, the chamfer 30 at an angle of about 23 ° to the longitudinal axis.
  • At least one sealing element 8a, 8b is formed as a brush seal. It should also be mentioned that the sealing arrangement 6 is not limited to the application described here in the low-pressure turbine area. The sealing arrangement 6, in addition to the illustrated outer sealing arrangement as inner
  • a sealing arrangement 6 for a gas turbine in particular for an aircraft gas turbine 1, with at least one sealing element 8a, 8b, the sealing element 8a, 8b being arranged in an annular space 12 between a housing 14 and at least one row of blades 4, in particular a row of blades wherein the annular space 12 is traversed by a leakage flow 16, which is branched off from a main stream 18 and flows back to the annular space 12 at a reentry point 20 in the main flow 18, wherein the at least one sealing element 8b associated with the housing 14 and in the direction of Blade row 4 extends and a leakage point 22 between the sealing element 8b and the blade row 4 in the region of the reentry point 20 of the leakage flow 16 is arranged in the main flow 18.
  • a gas turbine in particular an aircraft gas turbine 1, with such a sealing arrangement 6.

Abstract

L'invention concerne un ensemble d'étanchéité (6) pour une turbine à gaz, notamment une turbine à gaz d'avion (1), comportant au moins un élément d'étanchéité (8a,8b) qui est disposé dans un espace annulaire (12) entre un carter (14) et au moins une rangée d'aubes (4), notamment une rangée d'aubes mobiles, l'espace annulaire (12) étant traversé par un écoulement de fuite (16) dérivant d'un écoulement principal (18) auquel il retourne en un point de retour (20), après avoir traversé l'espace annulaire (12). Selon l'invention, l'élément d'étanchéité (8b) est associé au carter (14) et s'étend dans la direction de la rangée d'aubes (4) et un point de fuite (22) est disposé entre l'élément d'étanchéité (8b) et la rangée d'aubes (4), dans la zone du point de retour (20) de l'écoulement de fuite (16) dans l'écoulement principal (18). L'invention porte également sur une turbine à gaz, notamment une turbine à gaz d'avion (1), dotée d'un ensemble d'étanchéité (6) de ce type.
PCT/DE2010/001279 2009-11-07 2010-10-30 Ensemble d'étanchéité pour une turbine à gaz et turbine à gaz de ce type WO2011054341A2 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102009052314.6 2009-11-07
DE200910052314 DE102009052314A1 (de) 2009-11-07 2009-11-07 Dichtanordnung für eine Gasturbine und eine derartige Gasturbine

Publications (2)

Publication Number Publication Date
WO2011054341A2 true WO2011054341A2 (fr) 2011-05-12
WO2011054341A3 WO2011054341A3 (fr) 2011-07-07

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/DE2010/001279 WO2011054341A2 (fr) 2009-11-07 2010-10-30 Ensemble d'étanchéité pour une turbine à gaz et turbine à gaz de ce type

Country Status (2)

Country Link
DE (1) DE102009052314A1 (fr)
WO (1) WO2011054341A2 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2647796A1 (fr) * 2012-04-04 2013-10-09 MTU Aero Engines GmbH Système d'étanchéité pour turbomachine
EP2647795A1 (fr) * 2012-04-04 2013-10-09 MTU Aero Engines GmbH Système d'étanchéité pour turbomachine
EP2657452A1 (fr) * 2010-12-22 2013-10-30 Mitsubishi Heavy Industries, Ltd. Turbine
WO2013162946A1 (fr) 2012-04-24 2013-10-31 United Technologies Corporation Aube à élément poreux abradable
EP2623722A4 (fr) * 2010-09-28 2017-12-13 Mitsubishi Hitachi Power Systems, Ltd. Turbine

Citations (1)

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Publication number Priority date Publication date Assignee Title
DE19807247A1 (de) 1998-02-20 1999-09-09 Mtu Muenchen Gmbh Strömungsmaschine mit Rotor und Stator

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EP0894945A3 (fr) * 1997-07-29 2000-07-12 Siemens Aktiengesellschaft Turbine et aubage de turbine
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EP0982475A1 (fr) * 1998-08-28 2000-03-01 Asea Brown Boveri AG Turbomachine à écoulement axial avec aubage caréné
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EP1001139B1 (fr) * 1998-11-10 2004-01-07 ALSTOM (Switzerland) Ltd Dispositif d'étanchéité pour les extrémités des aubes de turbine
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US6896482B2 (en) * 2003-09-03 2005-05-24 General Electric Company Expanding sealing strips for steam turbines
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EP2096262A1 (fr) * 2008-02-26 2009-09-02 Siemens Aktiengesellschaft Turbine à flux axial avec des pertes de fuites réduites

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DE19807247A1 (de) 1998-02-20 1999-09-09 Mtu Muenchen Gmbh Strömungsmaschine mit Rotor und Stator

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2623722A4 (fr) * 2010-09-28 2017-12-13 Mitsubishi Hitachi Power Systems, Ltd. Turbine
EP2657452A1 (fr) * 2010-12-22 2013-10-30 Mitsubishi Heavy Industries, Ltd. Turbine
EP2657452A4 (fr) * 2010-12-22 2014-06-11 Mitsubishi Heavy Ind Ltd Turbine
US9353640B2 (en) 2010-12-22 2016-05-31 Mitsubishi Hitachi Power Systems, Ltd. Turbine
EP2647796A1 (fr) * 2012-04-04 2013-10-09 MTU Aero Engines GmbH Système d'étanchéité pour turbomachine
EP2647795A1 (fr) * 2012-04-04 2013-10-09 MTU Aero Engines GmbH Système d'étanchéité pour turbomachine
US20130266427A1 (en) * 2012-04-04 2013-10-10 Mtu Aero Engines Gmbh Sealing system for a turbomachine
WO2013162946A1 (fr) 2012-04-24 2013-10-31 United Technologies Corporation Aube à élément poreux abradable
EP2841696A4 (fr) * 2012-04-24 2016-06-08 United Technologies Corp Aube à élément poreux abradable
US9879559B2 (en) 2012-04-24 2018-01-30 United Technologies Corporation Airfoils having porous abradable elements
EP3617451A1 (fr) * 2012-04-24 2020-03-04 United Technologies Corporation Aube à élément poreux abradable et procédé d'usinage associé

Also Published As

Publication number Publication date
WO2011054341A3 (fr) 2011-07-07
DE102009052314A1 (de) 2011-05-12

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