EP2504529B1 - Isolation d'un rebord circonférentiel d'un carter externe de turbomachine vis-à-vis d'un secteur d'anneau correspondant - Google Patents

Isolation d'un rebord circonférentiel d'un carter externe de turbomachine vis-à-vis d'un secteur d'anneau correspondant Download PDF

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Publication number
EP2504529B1
EP2504529B1 EP10805261.4A EP10805261A EP2504529B1 EP 2504529 B1 EP2504529 B1 EP 2504529B1 EP 10805261 A EP10805261 A EP 10805261A EP 2504529 B1 EP2504529 B1 EP 2504529B1
Authority
EP
European Patent Office
Prior art keywords
annular
bottom wall
outer casing
ring sector
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.)
Active
Application number
EP10805261.4A
Other languages
German (de)
English (en)
French (fr)
Other versions
EP2504529A1 (fr
Inventor
Fabrice Marcel Noël GARIN
Alain Dominique Gendraud
Gilles Jeannin
Sébastien Jean Laurent Prestel
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.)
Safran Aircraft Engines SAS
Original Assignee
SNECMA Services SA
SNECMA SAS
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Filing date
Publication date
Application filed by SNECMA Services SA, SNECMA SAS filed Critical SNECMA Services SA
Publication of EP2504529A1 publication Critical patent/EP2504529A1/fr
Application granted granted Critical
Publication of EP2504529B1 publication Critical patent/EP2504529B1/fr
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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
    • 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
    • F01D25/246—Fastening of diaphragms or stator-rings
    • 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

Definitions

  • the present invention relates to a turbine stage of a turbomachine, such as an airplane turbojet or turboprop.
  • a turbomachine low-pressure turbine comprises several stages each comprising a distributor formed of an annular row of stationary vanes carried by an outer casing and of a rotor wheel rotatably mounted downstream of the distributor in a cylindrical or frustoconical envelope formed by ring sectors circumferentially attached end to end on the outer casing.
  • Pressurized hot gases leaving the combustion chamber of the turbomachine pass between the vanes of the distributors and flow on the blades of the turbine wheels, which has the effect of raising the temperature of the envelopes formed by the sectors of the turbine. 'ring.
  • the outer casing has at least one circumferential hooking flange of the downstream ends of the ring sectors.
  • each ring sector has a downstream end formed with an annular cavity delimited by an upstream annular abutment, a downstream annular abutment and a bottom wall and this cavity is engaged on the circumferential rim of the casing, the sector ring being held in axial position on the flange by the annular abutments of the cavity.
  • the invention aims in particular to provide a simple, effective and economical solution to this problem.
  • a turbine stage of a turbomachine comprising a rotor wheel mounted inside a sectored ring carried by an outer casing, each ring sector having a downstream end formed with an annular cavity. defined by an upstream annular abutment, a downstream annular abutment and a bottom wall, the outer casing having at least one circumferential flange housed in this cavity for attachment of the downstream end of the ring sector, the bottom wall of the the annular cavity of the ring sector remaining spaced radially from the circumferential rim of the outer casing, so as to provide a thermally insulating space between them and comprises radial positioning means on the circumferential rim.
  • the radial positioning means comprise at least two studs formed projecting on the bottom wall of the annular cavity.
  • the contact surface between the ring sector and the circumferential flange is thus limited to the surface of the end of the pads.
  • the studs are located at the circumferential ends of the bottom wall.
  • the pads are located at a distance from the median axial plane of the bottom wall, so as to ensure a good radial positioning of the ring sector.
  • the pads are located between the median axial plane and the circumferential ends of the bottom wall, so as to limit the wear of the aforementioned elements in contact.
  • each annular abutment has a radial surface extending over the entire circumference of the ring sector, the circumferential rim of the outer casing being mounted without clearance between these surfaces of the annular abutments of the ring sector.
  • the pads may be parallelepipedic.
  • circumferential rim of the outer casing is also advantageous for the circumferential rim of the outer casing to be axially constrained between the annular abutments so as to guarantee the correct positioning of the ring sector on the outer casing.
  • the ratio between the contact surface of the pads and the bottom surface of the annular cavity is between 0.1 and 0.25.
  • the invention further relates to a turbomachine such as a turbojet or an airplane turboprop, characterized in that it comprises a low pressure turbine stage according to the invention.
  • the Figures 1 to 3 represent a turbomachine low-pressure turbine 1 of the prior art, comprising a plurality of stages each comprising a stationary nozzle distributor 2 carried by an outer casing 4 of the turbine, and a rotor wheel 5 mounted downstream of the distributor 2 and rotating in a substantially frustoconical envelope formed by ring sectors 6 carried circumferentially end to end by the casing 4 of the turbine.
  • the distributors 2 comprise walls of internal (not visible) and external revolution 7 which delimit between them an annular stream 8 of gas flow in the turbine and which are radially connected by the blades 3.
  • the rotor wheels 2 are integral with a turbine shaft, not shown, and each comprise outer and inner rings 9 (not visible), the outer ring 9 comprising external radial ribs 10 externally surrounded with a small clearance by the sectors ring 6.
  • Each ring sector 6 comprises a frustoconical wall 11 and a block of abradable material 12 fixed by soldering and / or welding on the radially inner surface of the frustoconical wall 11, this block 12 being of the honeycomb type and being intended to wear by friction on the ribs 10 of the wheel 5 to minimize the radial clearances between the wheel 5 and the ring sectors 6.
  • the frustoconical wall 11 of the ring sector has a downstream end 13 formed with an annular cavity open towards the outside and delimited by an upstream annular abutment 14, a downstream annular abutment 15 and a bottom wall 16.
  • Each annular abutment 14, 15 comprises a surface extending over the entire circumference of the ring sector 6.
  • the bottom wall 16 further has a downstream annular groove 17 and an upstream annular groove 18, allowing the cavity to be machined (see FIG. figure 3 ).
  • each ring sector 6 is engaged in an annular space 19 delimited between two annular flanges of the outer wall 7 of the downstream distributor 2, respectively a radially inner flange 20 and a radially outer flange 21, oriented upstream.
  • the outer casing 4 comprises an inner circumferential rim 22 whose cross section has the shape of a hook facing downstream, engaged in the cavity of the frustoconical wall 11 of the ring sector and held therein by the rim radially. external 21 of the distributor 2.
  • the circumferential flange 22 of the outer casing 4 is axially constrained, between the annular abutments 14, 15 of the ring sector 6, this stress remaining in all phases of operation of the turbomachine.
  • said flange 22 having a radially outer annular surface bearing against the radially outer flange 21 of the distributor and a radially inner annular surface bearing against the bottom wall 16 of the ring sector.
  • An axial clearance j1 is provided between the upstream end of the radially outer flange 21 and the connection zone 23 between the flange 22 and the outer casing 4. This clearance makes it possible to compensate for the effects of expansion and can become almost zero during the operation of the flange. the turbomachine.
  • the ring sector 6 is thus locked, at its downstream end 13, on the circumferential rim 22 of the housing by the distributor 2, the seal between the circumferential flange 22 and the ring sector 6 being provided by the axial stops 14 , And by the bottom wall 16.
  • the ring sector 6 is also hooked at its upstream end on the housing by means whose structure will not be detailed here.
  • the gases from the combustion chamber heat the ring sectors 6, the heat then being transmitted by conduction to the circumferential rim 22 of the housing.
  • the conduction surface or contact surface between the ring sector 6 and the circumferential rim 22 is important, so that in practice the temperature of the rim 22 can reach a limit value, for example 730.degree. is the maximum acceptable for the material conventionally used.
  • FIGS. Figures 4 to 6 A ring sector according to the invention is illustrated in FIGS. Figures 4 to 6 . It differs from that described above in that the bottom wall 16 of the annular cavity comprises at least two studs 24 projecting radially outwardly, the ends of which form bearing surfaces 25 on the circumferential edge 22.
  • the pads 24 are preferably arranged near the upstream stop 14 of the ring sector 6.
  • the ratio between the contact surface of the pads 24 and the bottom surface 16 is between 0.1 and 0.25.
  • such a structure reduces by approximately 40 ° C the temperature of the circumferential rim 22 during operation of the turbomachine.
  • the studs 24 are of parallelepipedal shape and located at the circumferential ends of the bottom wall 16.
  • the studs 24 are located at a distance from a median axial plane P of the bottom wall 16, on either side of it and are located between the median axial plane P and one of the circumferential ends of the bottom wall 16.
  • each ring sector is immobilized circumferentially on the casing by means located in its median plane P, it expands relative to to the housing on either side of the median plane P.
  • the bringing of the pads 24 of the plane P thus reduces their friction on the circumferential rim 22 of the housing.
  • Positioning them at a distance from the plane P ensures a good radial positioning of the ring sector on the circumferential flange 22, avoiding any risk of tilting of the ring sector on one side or the other of the median plane P.
  • the studs 24 could have any other desired shape, for example square, cylindrical, frustoconical, etc.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Supercharger (AREA)
EP10805261.4A 2009-11-25 2010-11-24 Isolation d'un rebord circonférentiel d'un carter externe de turbomachine vis-à-vis d'un secteur d'anneau correspondant Active EP2504529B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0905657A FR2952965B1 (fr) 2009-11-25 2009-11-25 Isolation d'un rebord circonferentiel d'un carter externe de turbomachine vis-a-vis d'un secteur d'anneau correspondant
PCT/FR2010/052495 WO2011064496A1 (fr) 2009-11-25 2010-11-24 Isolation d'un rebord circonférentiel d'un carter externe de turbomachine vis-à-vis d'un secteur d'anneau correspondant

Publications (2)

Publication Number Publication Date
EP2504529A1 EP2504529A1 (fr) 2012-10-03
EP2504529B1 true EP2504529B1 (fr) 2013-10-09

Family

ID=42312955

Family Applications (1)

Application Number Title Priority Date Filing Date
EP10805261.4A Active EP2504529B1 (fr) 2009-11-25 2010-11-24 Isolation d'un rebord circonférentiel d'un carter externe de turbomachine vis-à-vis d'un secteur d'anneau correspondant

Country Status (9)

Country Link
US (1) US8961117B2 (ru)
EP (1) EP2504529B1 (ru)
JP (1) JP5771217B2 (ru)
CN (1) CN102630268B (ru)
BR (1) BR112012012393B1 (ru)
CA (1) CA2781936C (ru)
FR (1) FR2952965B1 (ru)
RU (1) RU2548535C2 (ru)
WO (1) WO2011064496A1 (ru)

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WO2013163581A1 (en) * 2012-04-27 2013-10-31 General Electric Company System and method of limiting axial movement between a hanger and a fairing assembly in a turbine assembly
EP2696037B1 (de) * 2012-08-09 2017-03-01 MTU Aero Engines AG Abdichtung des Strömungskanals einer Strömungsmaschine
JP6233578B2 (ja) * 2013-12-05 2017-11-22 株式会社Ihi タービン
US10655495B2 (en) 2017-02-24 2020-05-19 General Electric Company Spline for a turbine engine
US10648362B2 (en) 2017-02-24 2020-05-12 General Electric Company Spline for a turbine engine
US20180347399A1 (en) * 2017-06-01 2018-12-06 Pratt & Whitney Canada Corp. Turbine shroud with integrated heat shield
FR3071273B1 (fr) * 2017-09-21 2019-08-30 Safran Aircraft Engines Ensemble d'etancheite de turbine pour turbomachine
US10982559B2 (en) * 2018-08-24 2021-04-20 General Electric Company Spline seal with cooling features for turbine engines
FR3096395B1 (fr) * 2019-05-21 2021-04-23 Safran Aircraft Engines Turbine pour une turbomachine, telle qu’un turboréacteur ou un turbopropulseur d’avion
FR3100838B1 (fr) * 2019-09-13 2021-10-01 Safran Aircraft Engines Anneau d’etancheite de turbomachine
FR3109402B1 (fr) * 2020-04-15 2022-07-15 Safran Aircraft Engines Turbine pour une turbomachine

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US4687413A (en) * 1985-07-31 1987-08-18 United Technologies Corporation Gas turbine engine assembly
FR2635562B1 (fr) * 1988-08-18 1993-12-24 Snecma Anneau de stator de turbine associe a un support de liaison au carter de turbine
US5553999A (en) * 1995-06-06 1996-09-10 General Electric Company Sealable turbine shroud hanger
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FR2800797B1 (fr) * 1999-11-10 2001-12-07 Snecma Assemblage d'un anneau bordant une turbine a la structure de turbine
DE10122464C1 (de) * 2001-05-09 2002-03-07 Mtu Aero Engines Gmbh Mantelring
US6733235B2 (en) * 2002-03-28 2004-05-11 General Electric Company Shroud segment and assembly for a turbine engine
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US20040219011A1 (en) * 2003-05-02 2004-11-04 General Electric Company High pressure turbine elastic clearance control system and method
JP4200846B2 (ja) * 2003-07-04 2008-12-24 株式会社Ｉｈｉ シュラウドセグメント
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Also Published As

Publication number Publication date
CA2781936A1 (fr) 2011-06-03
FR2952965A1 (fr) 2011-05-27
JP2013512382A (ja) 2013-04-11
CN102630268A (zh) 2012-08-08
BR112012012393A2 (pt) 2016-04-12
CN102630268B (zh) 2015-07-08
WO2011064496A1 (fr) 2011-06-03
RU2548535C2 (ru) 2015-04-20
US8961117B2 (en) 2015-02-24
CA2781936C (fr) 2017-12-12
JP5771217B2 (ja) 2015-08-26
US20120288362A1 (en) 2012-11-15
RU2012126095A (ru) 2013-12-27
FR2952965B1 (fr) 2012-03-09
EP2504529A1 (fr) 2012-10-03
BR112012012393B1 (pt) 2020-11-10

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