EP1586743B1 - Anneau de turbine - Google Patents

Anneau de turbine Download PDF

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Publication number
EP1586743B1
EP1586743B1 EP05290821A EP05290821A EP1586743B1 EP 1586743 B1 EP1586743 B1 EP 1586743B1 EP 05290821 A EP05290821 A EP 05290821A EP 05290821 A EP05290821 A EP 05290821A EP 1586743 B1 EP1586743 B1 EP 1586743B1
Authority
EP
European Patent Office
Prior art keywords
ring according
sectors
tongues
slots
tongue
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
EP05290821A
Other languages
German (de)
English (en)
French (fr)
Other versions
EP1586743A1 (fr
Inventor
Ludovic Nicollas
Nicolas Hervy
Marc Roger Marchi
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 SAS
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 SNECMA SAS filed Critical SNECMA SAS
Publication of EP1586743A1 publication Critical patent/EP1586743A1/fr
Application granted granted Critical
Publication of EP1586743B1 publication Critical patent/EP1586743B1/fr
Active legal-status Critical Current
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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
    • 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

Definitions

  • the invention relates to a turbine ring forming the outer casing of the rotor of this turbine.
  • the invention is particularly applicable to a high pressure turbine located immediately downstream of the combustion chamber of an aircraft turbojet engine. It relates more particularly to the connection and cooling of the sectors constituting said turbine ring.
  • the rotor rotates inside a fixed turbine ring formed by a plurality of curved sectors joined end to end, circumferentially, to form the envelope. rotor.
  • the temperature of the gases driving the impeller is such that the thermomechanical stresses that are created in and between the sectors can cause deterioration reducing the life of the rings.
  • the formation of small cracks and / or spalling of the inner (so-called hot surface) side of the sectors is often observed, mainly in the vicinity of the connections between adjacent sectors.
  • sealing systems are provided between these adjacent sectors, comprising tabs extending between these sectors. and housed in slots made vis-à-vis in the adjacent radial faces of said sectors.
  • a known sector 1 represented in figure 1 comprises a sealing system comprising four tongues 2-5, housed in slots 6, 7, 8.
  • the tongue 3 is folded and extends between two slots 6, 7 opening into one another and welcoming d Other straight tabs 2, 4.
  • the slots are difficult to machine accurately, especially because of the difference in thickness required to insert the folded tab. The positioning of the latter is delicate.
  • the tongue 2 is entirely housed in a slot 6 parallel to the hot face 9 of the sector and at a short distance therefrom.
  • the practice of the slots creates areas of stress concentration which, when they are close to the surface hot, weaken the room and accelerate its deterioration.
  • the invention makes it possible to eliminate these disadvantages.
  • the invention relates to a turbine ring forming a rotor casing, of the type constituted by a plurality of sectors, joined end to end with the interposition of sealing systems comprising tabs extending between adjacent sectors, said tabs being accommodated. in slots made opposite each other in adjacent radial faces of said sectors, characterized in that each sealing system is constituted by straight tabs engaged in respective rectilinear slots of said radial faces and that the slots made on each radial face are independent, that is to say that said slots do not communicate with each other.
  • each sealing system comprises a first and a second tongues extending chevron on the inside of said radial faces, said tabs being engaged in slots rectilinear of said radial faces defining their relative positions with precision.
  • the air leak between two consecutive sectors can be precisely calibrated. This leak can therefore be identical in all inter-sector spaces.
  • the leak rate can be reduced by 10 to 20% compared to the configuration of the prior art described above.
  • Another advantage of the invention lies in the fact that the arrangement of the chevron tongues on the side of the hot face makes it possible both to move the stress concentration zones away from said hot face (since the slits deviate from the latter) and also to provide sufficient space between the tabs and the hot face to open air ejection channels cooling supplied from a cavity in the sector itself.
  • each sector comprises a cooling air circulation cavity, characterized in that it further comprises air ejection channels s extending between said cavity and at least one radial face of the sector, these channels opening on said radial face between an inner edge thereof and said first and second tongues.
  • turbine ring sectors 11 constituting the fixed casing of a rotor, not shown.
  • This is the high-pressure turbine of a turbojet engine.
  • This turbine is placed downstream of the combustion chamber.
  • such a ring consists of 32 ring 11 sectors curves such as those shown end to end to form a slightly conical envelope surrounding said rotor.
  • Each sector 11 is constituted a thick, slightly curved plate to reconstitute the ring.
  • Each sector 11 further comprises two radial faces 20, 21 through which it connects circumferentially to neighboring sectors, by sealing systems 26 (see FIG. figure 2 ) mentioned above.
  • Each sealing system 26 consists of a set of tabs engaged in corresponding slots defined in said radial faces 20, 21 facing each other.
  • Each tongue is engaged in two slots belonging to two ring sectors circumferentially adjacent.
  • each sector 11 is hollow and comprises a cooling air circulation cavity 35 fed from the outside.
  • the figure 4 illustrates very schematically the position of the ring formed by all of the sectors 11.
  • a turbine casing 15 defines with this ring an annular cavity 17.
  • the assembly extends radially outside the impeller with high pressure 19 itself axially interposed between the high pressure distributor 21 and the low pressure distributor 23. Air from the compressor is taken upstream of the combustion chamber and enters (via holes) into the annular cavity 17. This cavity fed so all ring sectors.
  • Each ring sector ( figure 3 ) comprises two distinct cavities 39 and 40 in the shape of a paper clip, separated by a partition 42 and respectively fed by orifices 37 and 38.
  • the air flowing in the cavity 39 escapes through a series of channels of objection 44 opening on the inlet edge 16 of the ring sector while the air flowing in the cavity 40 escapes through a series of ejection channels 46 opening on the output edge 18 of the ring sector.
  • the invention relates in particular to an advantageous development of said inter-sector sealing systems.
  • each sealing system 26 here consists of three straight tabs engaged in respective straight slots of the radial faces of the two adjacent sectors.
  • each sealing system ( figure 2 ) has a first tongue 27 and a second tongue 28, located on the inside of said radial faces, that is to say on the hot faces side of the sectors.
  • the tongues 27, 28 are arranged in chevron, that is to say engaged in slots 31, 32 of said radial faces which extend obliquely with respect to the inner faces 12 and outer 14 of the sectors. These slots define the relative positions of the two tabs.
  • each sealing system comprises a third tongue 29 extending substantially from one end to the other of the adjacent sectors, parallel to the axis of the ring, on the outer side of said radial faces.
  • the tongue 29 is engaged in rectilinear slots 33 of adjacent sectors.
  • the first tongue 27 extends between a point A located near the inlet edge of the two sectors, towards the inside (near the hot faces) and a point B located near the third tongue 29.
  • the second tongue 28 is positioned so that it extends between a point C located near the outlet edge 18 of each of the two sectors, inwardly and a point D located near the first tongue, substantially between the middle and two-thirds of it from point A.
  • the length of the first tongue 27 depends on the angle it makes with the third tongue 29. Once this angle is determined (several possibilities are represented on the figure 5 ) the position and the length of the second tab derive from it.
  • the angle defined by the first and third tab can be between 15 and 70 °, approximately.
  • the slots can be machined precisely and are perfectly localized.
  • the tabs can be inserted into these slots and their relative positions can be perfectly controlled. It follows that the leakage section between said first and second tongues (S 1 ) and the leakage section between the first and third tongues (S 2 ) are perfectly controlled.
  • each sector comprises air ejection channels. 50, extending between the cavity 40 and at least one radial face of the sector. These channels open on the radial face 20 between the inner edge thereof (hot face) and said first and second tongues 27, 28.
  • the chevron arrangement of the two tongues makes it possible to practice these air ejection channels.
  • the channels are arranged in a row parallel to the axis of the ring. In the example of the figure 3 they all extend perpendicularly at the radial face.
  • some channels 50 extend perpendicular to the radial face but others at the ends of said row or at least one of them are made at an angle and diverging with respect to the first, here in a direction from the cavity towards the radial face.
  • the angle between the diverging channels may be between 10 and 120 °
  • the parallel channels are at an angle to a direction perpendicular to the radial face. The angle is such that the air is ejected obliquely towards the rear of the ring.
  • the parallel channels are at an angle to a direction perpendicular to the radial face. The angle is such that the air is ejected obliquely towards the front of the ring.
  • the channels 50 open on the radial face 20 which is that the blades first reach given the direction of rotation indicated by the arrow F. This is favorable to avoid or limit reintroductions of hot gas in the spaces inter-segment.
  • the air escaping channels 50 cools the wall in which they are practiced by convection (thermal pumping) while the opposite wall (face 21) is cooled by the impact of the air jets.
  • the air jets escaping from the ducts 50 establish a kind of fluidic system preventing the ingestion of hot gases.
  • the slots 31, 32, 33 are independent, that is to say they do not communicate with each other. This avoids having to make remains at the junction of two slots. Inter-sector leakage sections are also reduced.
  • the invention also relates to any ring sector or ring sector assembly having the characteristics described above.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP05290821A 2004-04-15 2005-04-14 Anneau de turbine Active EP1586743B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0403925 2004-04-15
FR0403925A FR2869070B1 (fr) 2004-04-15 2004-04-15 Anneau de turbine

Publications (2)

Publication Number Publication Date
EP1586743A1 EP1586743A1 (fr) 2005-10-19
EP1586743B1 true EP1586743B1 (fr) 2012-05-30

Family

ID=34942125

Family Applications (1)

Application Number Title Priority Date Filing Date
EP05290821A Active EP1586743B1 (fr) 2004-04-15 2005-04-14 Anneau de turbine

Country Status (9)

Country Link
US (1) US7513740B1 (uk)
EP (1) EP1586743B1 (uk)
JP (1) JP4679215B2 (uk)
CN (1) CN1683772B (uk)
CA (1) CA2503066C (uk)
ES (1) ES2386146T3 (uk)
FR (1) FR2869070B1 (uk)
RU (1) RU2377419C2 (uk)
UA (1) UA91958C2 (uk)

Families Citing this family (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1991762B1 (de) * 2006-03-06 2015-03-11 Alstom Technology Ltd Gasturbine mit ringförmigem hitzeschild und abgewinkelten dichtungsstreifen
FR2919345B1 (fr) * 2007-07-26 2013-08-30 Snecma Anneau pour une roue de turbine de turbomachine.
US7874792B2 (en) 2007-10-01 2011-01-25 United Technologies Corporation Blade outer air seals, cores, and manufacture methods
US8075255B2 (en) * 2009-03-31 2011-12-13 General Electric Company Reducing inter-seal gap in gas turbine
US20130134678A1 (en) * 2011-11-29 2013-05-30 General Electric Company Shim seal assemblies and assembly methods for stationary components of rotary machines
US9810086B2 (en) * 2011-11-06 2017-11-07 General Electric Company Asymmetric radial spline seal for a gas turbine engine
US9863323B2 (en) 2015-02-17 2018-01-09 General Electric Company Tapered gas turbine segment seals
US10689994B2 (en) * 2016-03-31 2020-06-23 General Electric Company Seal assembly to seal corner leaks in gas turbine
US20180355754A1 (en) * 2017-02-24 2018-12-13 General Electric Company Spline for a turbine engine
US20180355741A1 (en) * 2017-02-24 2018-12-13 General Electric Company Spline for a turbine engine
US10648362B2 (en) * 2017-02-24 2020-05-12 General Electric Company Spline for a turbine engine
US20180340437A1 (en) * 2017-02-24 2018-11-29 General Electric Company Spline for a turbine engine
US10655495B2 (en) * 2017-02-24 2020-05-19 General Electric Company Spline for a turbine engine
FR3070718B1 (fr) * 2017-09-06 2019-08-23 Safran Aircraft Engines Ensemble de turbine a secteurs d'anneau
US10982559B2 (en) * 2018-08-24 2021-04-20 General Electric Company Spline seal with cooling features for turbine engines

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2597921A1 (fr) * 1986-04-24 1987-10-30 Snecma Anneau de turbine sectorise
US5127793A (en) * 1990-05-31 1992-07-07 General Electric Company Turbine shroud clearance control assembly
US5088888A (en) * 1990-12-03 1992-02-18 General Electric Company Shroud seal
FR2758856B1 (fr) * 1997-01-30 1999-02-26 Snecma Joint d'etancheite a plaquettes empilees glissant dans des fentes de reception
FR2800797B1 (fr) * 1999-11-10 2001-12-07 Snecma Assemblage d'un anneau bordant une turbine a la structure de turbine
US6340285B1 (en) * 2000-06-08 2002-01-22 General Electric Company End rail cooling for combined high and low pressure turbine shroud
US6814538B2 (en) * 2003-01-22 2004-11-09 General Electric Company Turbine stage one shroud configuration and method for service enhancement

Also Published As

Publication number Publication date
FR2869070A1 (fr) 2005-10-21
JP2005299663A (ja) 2005-10-27
UA91958C2 (uk) 2010-09-27
US20090074579A1 (en) 2009-03-19
EP1586743A1 (fr) 2005-10-19
CN1683772A (zh) 2005-10-19
FR2869070B1 (fr) 2008-10-17
CN1683772B (zh) 2011-07-06
RU2005110997A (ru) 2006-10-20
CA2503066A1 (fr) 2005-10-15
RU2377419C2 (ru) 2009-12-27
ES2386146T3 (es) 2012-08-10
CA2503066C (fr) 2013-01-15
JP4679215B2 (ja) 2011-04-27
US7513740B1 (en) 2009-04-07

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