EP2696039A1 - Étage de turbine à gaz - Google Patents

Étage de turbine à gaz Download PDF

Info

Publication number
EP2696039A1
EP2696039A1 EP12179978.7A EP12179978A EP2696039A1 EP 2696039 A1 EP2696039 A1 EP 2696039A1 EP 12179978 A EP12179978 A EP 12179978A EP 2696039 A1 EP2696039 A1 EP 2696039A1
Authority
EP
European Patent Office
Prior art keywords
gas turbine
wall
circumferential groove
sealing ring
turbine stage
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.)
Granted
Application number
EP12179978.7A
Other languages
German (de)
English (en)
Other versions
EP2696039B1 (fr
Inventor
Hans-Peter Dr. Hackenberg
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 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 MTU Aero Engines AG filed Critical MTU Aero Engines AG
Priority to EP12179978.7A priority Critical patent/EP2696039B1/fr
Priority to ES12179978.7T priority patent/ES2547340T3/es
Priority to US13/959,021 priority patent/US20140044537A1/en
Publication of EP2696039A1 publication Critical patent/EP2696039A1/fr
Application granted granted Critical
Publication of EP2696039B1 publication Critical patent/EP2696039B1/fr
Not-in-force legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • 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
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/243—Flange connections; Bolting arrangements
    • 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
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved
    • F05D2250/711—Shape curved convex

Definitions

  • the present invention relates to a gas turbine stage, in particular a low-pressure compressor stage, with a vane assembly and a sealing ring element, a gas turbine, in particular an aircraft gas turbine engine with one or more such gas turbine stages, and a vane assembly and a sealing ring member for such a gas turbine.
  • an inner wall of one of a vane assembly and a sealing ring member engages a circumferential groove of the other of the vane assembly and the sealing ring member.
  • the inner wall has, according to in-house practice in a development in the circumferential direction or in a plan view of the spoke centering a rectangular cross section with two opposite, the Nutinnen vom the circumferential groove facing end faces and two against this by 90 ° angled flanks in one edge into each other pass.
  • the object of the present invention is to provide an improved gas turbine.
  • a gas turbine in particular an aircraft gas turbine engine, one or more gas turbine stages, in particular low-pressure compressor stages, which according to another aspect of the present invention each have a vane assembly with one or more vanes with a radially inner Leitschaufelfuß. Two or more vanes may be connected together in a development, in particular integrally.
  • a sealing ring element On the vane a sealing ring element is supported by a spokes centering, which has an inner wall and a circumferential groove receiving this.
  • a circumferential groove is understood as meaning, in particular, a recess or recess circulating in the circumferential direction and bounded on both sides by, in particular, parallel, groove inner surfaces.
  • a sealing ring element may be part of a multi-part closed sealing ring.
  • a one-piece sealing ring is generally referred to as a sealing ring element in the sense of the present invention.
  • the inner wall has one or more, in particular two opposite, preferably parallel end faces, each facing a Nutinnen dye the circumferential groove and in an undeformed state in particular, at least substantially, can be oriented normal to an axial direction of the gas turbine stage.
  • This adjacent, in particular connecting these, the inner wall one or more, in particular two opposing, preferably parallel flanks, which are each angled against the adjacent (s) end faces or form an angle with these, preferably greater than 75 ° and / or may be less than 105 °.
  • a rounding in particular a radius, is formed between at least one of these end faces and at least one of these flanks.
  • a rounding means in particular, a connecting surface which is mono- or multiply, in particular convexly curved, which may in particular have radially extending straight generatrices, wherein in one development the rounding can be formed kink-free or with continuous curvature.
  • the inner wall of the spoke centering can be connected to the guide blade root, in particular integrally, and the circumferential groove can be formed on the sealing ring element.
  • the inner wall with the sealing ring member, in particular integrally connected and the circumferential groove may be formed on the Leitschaufelfuß.
  • a vane assembly for the gas turbine (n-stage) described herein having an inner wall connected thereto for receiving in a circumferential groove to form a spoke centering of the seal member on the vane assembly, the inner wall at least one of a Nutinnen character the circumferential groove facing end face and one of these adjacent, against this angled edge, and wherein between the end face and the flank a rounding, in particular a radius is formed.
  • the present embodiments in particular advantageous developments, therefore relate equally to such a vane arrangement or such a sealing ring element, a gas turbine stage and a gas turbine with such a spoke centering.
  • the inner wall has at least one radial groove for receiving a sliding element, in particular a sliding block connected to the circumferential groove, for forming the spoke centering.
  • this radius has at least 25%, in particular at least 30%, of a width of the flank.
  • a width of a flank is understood in particular to mean its axial width or the wall thickness of the inner wall.
  • the radius may be at least 0.5 mm, in particular at least 10 mm, and preferably at least 15 mm.
  • a rounding can in particular together with the inner wall or its end face (s) and flank (s) urgeformt, in particular cast. Additionally or alternatively, the fillet may preferably be reshaped and / or machined or made.
  • a rounding can be hardened, in particular by a corresponding thermal method and / or a coating. As a result, in particular the wear in the area of the rounding can be further reduced.
  • the present invention is particularly advantageous for the spoke centering of a brush seal element.
  • a vane arrangement may have a torsion preferential direction, which may for example result from a setting and / or a profile of the guide vanes or the like.
  • this is understood in particular to mean a torsional direction about a radial axis into which the guide vane (s) of the vane arrangement are twisted during operation with increasing flow and / or power of the gas turbine (n-stage).
  • a rounding is then provided in particular also or only at one end of the end face (s) in the circumferential direction, which is delivered in torsion in the Torsionsvorzugsraum against the adjacent Nutinnen Diagram.
  • Fig. 1 shows a meridian section of a portion of a gas turbine with a gas turbine stage with a vane assembly according to an embodiment of the present invention.
  • Guide vane assembly shown in a perspective view has a plurality of integrally connected vanes 1 with a common, radially inner Leitschaufelfuß 1.1, from which a common inner wall 1.2 from a side facing away from the vane blade (bottom in Fig. 1 . 2 ) of the guide blade root 1.1 starting radially inward (down in Fig. 1 . 2 ).
  • a radial groove 1.3 is formed in the inner wall 1.2 a radial groove 1.3 is formed.
  • a sliding block is arranged radially movable, which is connected to a circumferential groove with two legs 2.1, 2.2, which is formed in a brush seal ring element 2.
  • the brush seal element 2 is mounted on the guide vane arrangement in a memory-centered manner.
  • the inner wall 1.2 which is disposed between the two legs 2.1, 2.2 of the circumferential groove of the brush seal member 2, two parallel planar axial end faces 1.4, which are oriented in an undeformed position of the vanes 3 normal to an axial direction of the gas turbine (n) stage ie in Fig. 3 run vertically.
  • These two end faces 1.4 are connected by two parallel planar flanks 1.5, which are angled perpendicular to the end faces 1.4.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP12179978.7A 2012-08-10 2012-08-10 Étage de turbine à gaz Not-in-force EP2696039B1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
EP12179978.7A EP2696039B1 (fr) 2012-08-10 2012-08-10 Étage de turbine à gaz
ES12179978.7T ES2547340T3 (es) 2012-08-10 2012-08-10 Grado de turbina de gas
US13/959,021 US20140044537A1 (en) 2012-08-10 2013-08-05 Gas turbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP12179978.7A EP2696039B1 (fr) 2012-08-10 2012-08-10 Étage de turbine à gaz

Publications (2)

Publication Number Publication Date
EP2696039A1 true EP2696039A1 (fr) 2014-02-12
EP2696039B1 EP2696039B1 (fr) 2015-07-29

Family

ID=46967917

Family Applications (1)

Application Number Title Priority Date Filing Date
EP12179978.7A Not-in-force EP2696039B1 (fr) 2012-08-10 2012-08-10 Étage de turbine à gaz

Country Status (3)

Country Link
US (1) US20140044537A1 (fr)
EP (1) EP2696039B1 (fr)
ES (1) ES2547340T3 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3208426A1 (fr) 2016-02-18 2017-08-23 MTU Aero Engines GmbH Segment d'aube directrice pour turbomachine
DE102016222608A1 (de) 2016-11-17 2018-05-17 MTU Aero Engines AG Dichtungsanordnung für eine Leitschaufelanordnung einer Gasturbine
EP3483399A1 (fr) * 2017-11-09 2019-05-15 MTU Aero Engines GmbH Dispositif d'étanchéité pour une turbomachine, procédé de fabrication d'un dispositif d'étanchéité et turbomachine
DE102018213983A1 (de) * 2018-08-20 2020-02-20 MTU Aero Engines AG Verstellbare Leitschaufelanordnung, Leitschaufel, Dichtungsträger und Turbomaschine
EP3885536A1 (fr) * 2020-03-25 2021-09-29 MTU Aero Engines AG Composant de turbine à gaz

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5224822A (en) * 1991-05-13 1993-07-06 General Electric Company Integral turbine nozzle support and discourager seal
US20030082050A1 (en) * 2001-10-29 2003-05-01 Emil Aschenbruck Device for sealing turbomachines
US20060062673A1 (en) * 2004-09-23 2006-03-23 Coign Robert W Mechanical solution for rail retention of turbine nozzles
GB2425155A (en) * 2005-04-13 2006-10-18 Rolls Royce Plc A mounting arrangement
US20070134088A1 (en) * 2005-12-08 2007-06-14 General Electric Company Methods and apparatus for assembling turbine engines
US20110150640A1 (en) * 2003-08-21 2011-06-23 Peter Tiemann Labyrinth Seal in a Stationary Gas Turbine
EP2360352A2 (fr) * 2010-02-12 2011-08-24 Rolls-Royce Deutschland Ltd & Co KG Joint d'étanchéité sans vis d'une turbine à gaz

Family Cites Families (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3829233A (en) * 1973-06-27 1974-08-13 Westinghouse Electric Corp Turbine diaphragm seal structure
US5749584A (en) * 1992-11-19 1998-05-12 General Electric Company Combined brush seal and labyrinth seal segment for rotary machines

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5224822A (en) * 1991-05-13 1993-07-06 General Electric Company Integral turbine nozzle support and discourager seal
US20030082050A1 (en) * 2001-10-29 2003-05-01 Emil Aschenbruck Device for sealing turbomachines
US20110150640A1 (en) * 2003-08-21 2011-06-23 Peter Tiemann Labyrinth Seal in a Stationary Gas Turbine
US20060062673A1 (en) * 2004-09-23 2006-03-23 Coign Robert W Mechanical solution for rail retention of turbine nozzles
GB2425155A (en) * 2005-04-13 2006-10-18 Rolls Royce Plc A mounting arrangement
US20070134088A1 (en) * 2005-12-08 2007-06-14 General Electric Company Methods and apparatus for assembling turbine engines
EP2360352A2 (fr) * 2010-02-12 2011-08-24 Rolls-Royce Deutschland Ltd & Co KG Joint d'étanchéité sans vis d'une turbine à gaz

Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10895162B2 (en) 2016-02-18 2021-01-19 MTU Aero Engines AG Guide vane segment for a turbomachine
DE102016202519A1 (de) 2016-02-18 2017-08-24 MTU Aero Engines AG Leitschaufelsegment für eine Strömungsmaschine
EP3208426B1 (fr) * 2016-02-18 2025-04-30 MTU Aero Engines AG Segment d'aube directrice pour turbomachine
EP3208426A1 (fr) 2016-02-18 2017-08-23 MTU Aero Engines GmbH Segment d'aube directrice pour turbomachine
EP3324001A1 (fr) * 2016-11-17 2018-05-23 MTU Aero Engines AG Agencement de joint d'étanchéité pour un ensemble d'aubes de guidage d'une turbine à gaz
US10677080B2 (en) 2016-11-17 2020-06-09 MTU Aero Engines AG Seal system for a guide blade system of a gas turbine
DE102016222608A1 (de) 2016-11-17 2018-05-17 MTU Aero Engines AG Dichtungsanordnung für eine Leitschaufelanordnung einer Gasturbine
US10865651B2 (en) 2017-11-09 2020-12-15 MTU Aero Engines AG Sealing assembly for a fluid kinetic machine, method for producing a sealing assembly as well as fluid kinetic machine
EP3483399A1 (fr) * 2017-11-09 2019-05-15 MTU Aero Engines GmbH Dispositif d'étanchéité pour une turbomachine, procédé de fabrication d'un dispositif d'étanchéité et turbomachine
DE102018213983A1 (de) * 2018-08-20 2020-02-20 MTU Aero Engines AG Verstellbare Leitschaufelanordnung, Leitschaufel, Dichtungsträger und Turbomaschine
EP3613952A1 (fr) 2018-08-20 2020-02-26 MTU Aero Engines GmbH Dispositif d'aube directrice réglable, aube directrice, porte-joint et turbomachine
US11300004B2 (en) 2018-08-20 2022-04-12 MTU Aero Engines AG Adjustable guide vane arrangement, guide vane, seal carrier and turbomachine
EP3885536A1 (fr) * 2020-03-25 2021-09-29 MTU Aero Engines AG Composant de turbine à gaz
DE102020203840A1 (de) 2020-03-25 2021-09-30 MTU Aero Engines AG Gasturbinenbauteil
US11585242B2 (en) 2020-03-25 2023-02-21 MTU Aero Engines AG Gas turbine component

Also Published As

Publication number Publication date
US20140044537A1 (en) 2014-02-13
EP2696039B1 (fr) 2015-07-29
ES2547340T3 (es) 2015-10-05

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