EP1445421B1 - Vorrichtung für die Belüftung eines Rotor einer Hochdruckturbine - Google Patents

Vorrichtung für die Belüftung eines Rotor einer Hochdruckturbine Download PDF

Info

Publication number
EP1445421B1
EP1445421B1 EP04100404A EP04100404A EP1445421B1 EP 1445421 B1 EP1445421 B1 EP 1445421B1 EP 04100404 A EP04100404 A EP 04100404A EP 04100404 A EP04100404 A EP 04100404A EP 1445421 B1 EP1445421 B1 EP 1445421B1
Authority
EP
European Patent Office
Prior art keywords
upstream
downstream
disk
turbine
flange
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.)
Expired - Lifetime
Application number
EP04100404A
Other languages
English (en)
French (fr)
Other versions
EP1445421A1 (de
Inventor
Patrick Rossi
Jean-Claude Christian Taillant
Maurice Guy Judet
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 EP1445421A1 publication Critical patent/EP1445421A1/de
Application granted granted Critical
Publication of EP1445421B1 publication Critical patent/EP1445421B1/de
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • F01D5/082—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades on the side of the rotor disc

Definitions

  • the present invention relates generally to the field of ventilation of a high-pressure turbine rotor of a turbomachine.
  • the invention relates to a ventilation device of a high pressure turbine rotor, comprising an upstream turbine disk and a downstream turbine disk.
  • FIG. 1 represents a conventional high-pressure turbine rotor 1 of the prior art, disposed downstream of a combustion chamber 2, and comprising an upstream turbine disk 3 equipped with vanes 4, as well as a disk downstream turbine 5 equipped with vanes 6.
  • the upstream disk 3 is provided on the one hand with an upstream flange 8 ensuring its fixing on a spacer 9 disposed around a rotor shaft 11 of a low-pressure turbine, and on the other hand with a downstream flange 10 fixedly assembled to an upstream flange 12 of the downstream disc 5.
  • an inter-disc seal 14, carried by a hollow structure 16 integral with a stationary distributor stage 18 or stator, is situated at the level of the assembly between the two flanges 10 and 12.
  • the inter-disc seal 14, of the labyrinth seal type thus makes it possible to create a separation between the two rotor stages 20 and 22, arranged on either side of the stage distributor 18.
  • downstream disk 5 comprises a downstream flange 13, also assembled on the spacer 9 surrounding the shaft 11 of the low pressure turbine.
  • a first cooling air flow D1 taken from the bottom of the combustion chamber 2 is delivered into a cavity 26 delimited on the one hand by means of a downstream face.
  • This air flow D1 is actually taken from the bottom of the chamber combustion 2, then conveyed in a cavity 30 in particular delimited by an upstream labyrinth seal 32 and a downstream labyrinth seal 34, via a conduit 28 disposed in an enclosure 29 separating the upstream labyrinth 24 from the bottom of the combustion chamber 2, and using injectors 36 arranged in the extension of the conduit 28 and opening into the cavity 30.
  • the seals 32 and 34 are arranged to be in contact with the upstream labyrinth 24 .
  • cooling air situated in the cavity 30 is able to penetrate into the cavity 26 by passing orifices 38 provided in an upstream part of the upstream labyrinth 24, these orifices 38 being axes substantially perpendicular to the longitudinal axis 40 of the turbine.
  • the cooling air flow D1 flows in the cavity 26 first longitudinally and then radially outwards along the upstream face of the upstream labyrinth 24 in order to cool it, then enters the cells 4a containing the feet of the blades 4 to cool them too.
  • a second cooling air flow D2 also taken from the bottom of the combustion chamber 2, enters the chamber 29 and flows through orifices 44 and 42 respectively provided in the upstream portion of the upstream labyrinth 24, and in the upstream flange 8 of the upstream disk 3.
  • the second cooling air flow D2 borrows an annular chamber 46 internally defined by the spacer 9, and externally delimited successively, from upstream to downstream, the flange 8, an inner bore 48 of the upstream disk 3, the flanges 10 and 12, an inner bore 50 of the downstream disk 5, and the flange 13.
  • a first portion D2a of the second cooling air flow D2 flows through orifices 52 formed in the downstream flange 10 of the upstream disk 3, in order to join the gap 19 located between the fixed distributor stage 18 and the rotor stage 20, as schematically represents the arrow referenced D2a.
  • the air flow of Diagrammatically shown in Figure 1 corresponds to an air leak at the cells 4a.
  • a second portion D2b of the second cooling air flow D2 flows through orifices 54 formed in the downstream flange 13 of the downstream disk 5, to penetrate inside a cavity 56 delimited by a part using an upstream face of a downstream labyrinth 58 disposed near the downstream disk 5, and secondly with the aid of a downstream face of the same downstream disk 5.
  • the second cooling air flow D2b circulates substantially radially in the cavity 56 outwards along the downstream face of the downstream labyrinth 58 in order to cool it, then enters cells 6a containing the roots of the vanes 6 so as to to cool them as well.
  • the rotor ventilation device thus has two separate cooling circuits, each associated with one of the two turbine disks, and respectively powered by the first and second air flow rates. D1 and D2 cooling.
  • the purpose of the invention is therefore to propose a device for ventilating a high-pressure turbine rotor of a turbomachine, the turbine being disposed downstream of a combustion chamber and comprising upstream and downstream turbine disks equipped with vane, the device comprising a cooling circuit provided with injectors arranged upstream of the upstream disk and being fed by a cooling air flow D taken from the bottom of the combustion chamber, the device at least partially overcoming the disadvantages mentioned herein above relative to the achievements of the prior art.
  • the subject of the invention is a device for ventilating a high-pressure turbine rotor of a turbomachine, the turbine being disposed downstream of a combustion chamber and comprising an upstream turbine disk equipped with blades and a downstream turbine disk also equipped with blades, the device comprising a cooling circuit provided with injectors arranged upstream of the upstream disk, the circuit being fed by a cooling air flow D taken from the bottom combustion chamber.
  • the cooling circuit is arranged in such a way that the flow of cooling air D coming from the injectors passes through orifices formed in an upstream flange of the an upstream disk allowing its attachment to an upstream flange of the downstream disk, so that this cooling air flow D flows axially downstream between an internal bore of the upstream disk and an upstream flange of the downstream disk allowing its attachment to a downstream flange a high-pressure compressor and the centering of the upstream disk, the ventilation device further comprising a single labyrinth integral with one of the two turbine disks and being interposed between these two disks, so that the air flow rate cooling circuit D separates into a first flow F1 flowing between a downstream face of the upstream disk and an upstream face of the single labyrinth in the direction of the blades of the upstream disk, and a second flow F2 flowing between an upstream face of the downstream disk and a face downstream of the single labyrinth towards the blades of the downstream disk.
  • the ventilation device no longer has two labyrinths respectively associated with upstream and downstream turbine disks, but has a single inter-disk labyrinth, each of the upstream and downstream faces is intended for guide a flow of cooling air towards the blades.
  • the reduction in the number of parts used consequently makes it possible to considerably reduce the mass, the bulk and the production cost of the rotor.
  • the specific positioning of the single labyrinth leads the latter to be less thermally stressed than a labyrinth arranged upstream of the upstream disk, mainly because of its location relative to the combustion chamber, and to the extent that the temperature of the cooling air flow D drops substantially as it passes through the inner bore of the upstream disk. This characteristic thus gives rise to an increase in the lifetime of this labyrinth, compared to the lifetime that an upstream labyrinth of the prior art could present.
  • the injection of the cooling air upstream of the upstream disk, the bypass of this upstream disk by the inner bore and the possibility of producing constituent elements of the rotor of small dimensions allows, by a single cavity defined jointly by a downstream face of the upstream disk and by an upstream face of the single labyrinth, to obtain sufficient pressure at the blades of the upstream disk.
  • the adjacent cavity delimited jointly by an upstream face of the downstream disk and by a downstream face of the single labyrinth is advantageously used to reduce the supply pressure of the vanes of the downstream disk.
  • the low pressure inside this adjacent cavity effectively avoids having to provide feed holes of the blades of too small dimensions, which are difficult to achieve.
  • the rotor made more compact by reducing the number of constituent elements of the rotor makes it possible to bring the under-chamber bearing of the upstream and downstream disks closer together, so that that it is then possible to obtain a better control of the games at the top of the blades, and therefore a better efficiency of the high pressure turbine.
  • FIG. 2 there is shown a turbine 100 at high pressure of a turbojet engine, comprising a device for ventilating the rotor of the turbine according to a preferred embodiment of the present invention.
  • a turbine 100 at high pressure of a turbojet engine comprising a device for ventilating the rotor of the turbine according to a preferred embodiment of the present invention.
  • the elements bearing the same reference numerals as those attached to the elements shown in Figure 1 correspond to the same or similar elements.
  • FIG. 2 shows a turbine 100 which differs first of all from the turbine 1 of the prior art in that a cooling air flow D, taken from the bottom of the combustion chamber 2 and adapted to passing through the injectors 36, is intended simultaneously to feed the vanes 4 and 6 of the upstream 3 and downstream 5 disks.
  • the cooling air from the combustion chamber 2 passes through the conduit 28 to join the injectors 36, this assembly consisting of the conduit 28 and the injectors 36 being located in a chamber 62 separating the upstream disk 3 bottom of the combustion chamber 2.
  • the cooling air flow D from the injectors 36 then enters a cavity 64 partially delimited by an upstream flange 66 of the upstream turbine disk 3, this upstream flange 66 whose main function is to ensure the attachment of this upstream disk 3 on an upstream flange 78 of the downstream disk 5.
  • this cavity 64 is also delimited jointly by the upstream seal 32 and the downstream seal 34, preferably of the labyrinth seal type, arranged near the injectors 36 respectively upstream and downstream of the latter.
  • the upstream gasket 32 cooperates with a downstream flange 70 of the high pressure turbine, this downstream flange 70 being arranged to be located radially outwardly relative to the upstream flange 66.
  • the upstream seal 32 closes the cavity 64 by marrying the upstream end of the upstream flange 66.
  • the downstream seal 34 cooperates with a secondary upstream flange 72 of the upstream turbine disk 3, arranged so as to be located radially towards the outside with respect to the upstream flange 66.
  • the cooling air escaping from the cavity 64 through the downstream gasket 34 can circulate radially outwards, along the upstream face of the upstream disk 3, by direction of blades 4.
  • Orifices 74 are provided in the upstream flange 66 of the upstream turbine disk 3 so that the flow of cooling air D can be conveyed towards the two turbine disks 3 and 5.
  • the orifices 74 are preferably arranged to be located radially next to the injectors 36.
  • the cooling air flow D enters an annular chamber 76 of axis 40, delimited externally via the upstream flange 66 of the upstream disk 3, and with the aid of the inner bore 48 of this same disc.
  • the annular chamber 76 is delimited internally by the upstream flange 78 of the downstream disc 5, this upstream flange 78 whose main function is to ensure the attachment of this downstream disk 5 to the upstream flange 66 of the upstream disk 3, and to center the assembly of the high-pressure turbine 100 on a downstream flange 79 d a high pressure compressor.
  • the cooling air flow D can then flow axially downstream between the inner bore 48 and the upstream flange 78, so that the upstream turbine disk 3 can be suitably cooled by contacting the cooling air with its internal bore 48.
  • the ventilation device comprises a single labyrinth 80 interposed between the turbine discs 3 and 5, and is integral with one of these two discs.
  • the single labyrinth 80 also called the inter-disk labyrinth, is attached to a secondary upstream flange 82 of the downstream turbine disk 5, the latter being arranged to be located radially outwardly. relative to the upstream flange 78.
  • the labyrinth 80 extends radially to match the fixed distributor stage 18 or stator provided between the two rotor stages 20 and 22, and has an inner bore 83 surrounding the upstream flange 78 of the disk 5, this bore 83 preferably having a diameter substantially identical to the diameter of the inner bore 48 of the disk 3.
  • the first stream F1 therefore flows in a cavity 68 located between the downstream face of the upstream turbine disk 3 and the upstream face of the labyrinth 80 in order to cool the downstream face of the disk 3, then enters cells 4a containing the feet of the blades 4. to cool them too.
  • the second flow F2 flows in a cavity 69 located between the upstream face of the downstream turbine disk 5 and the downstream face of the same labyrinth 80 to cool the upstream face of the disk 5, then enters the cells 6a containing the feet of the blades 6 to also cool them.
  • a plurality of orifices 84 is formed in the secondary upstream flange 82 of the downstream disk 5.
  • the ventilation device is such that the flow of cooling air D taken from the bottom of the combustion chamber 2 and intended to simultaneously feed the blades 4 and 6, borrows a single cooling circuit to the outlet of the passage between the bore 48 of the upstream disk 3 and the upstream flange 78 of the downstream turbine disk 5.
  • This specific characteristic greatly simplifies the design of the turbine 100 with respect to that of the turbine 1 of the prior art, in which two cooling air flows were taken at the bottom of the chamber of combustion 2, in order to borrow two totally separate cooling circuits.
  • the upstream flange 78 of the downstream turbine disk 5 comprises a plurality of orifices 86 able to be traversed by a third flow F3 of the cooling air flow D.
  • This third flow F3 is thus conveyed from the chamber annular 76 to an annular space 88 of the same axis, the space 88 being located between on the one hand the upstream flange 78 of the downstream disk 5 and the inner bore 50 of the same downstream disk 5, and on the other hand the spacer 9 disposed around the rotor shaft 11 of the low pressure turbine.
  • the cooling air flow F3 can flow axially downstream in the annular space 88, in order to cool the downstream disc 5 by contacting the air with its internal bore 50.
  • the third flow F3 is then evacuated downstream of the turbine 100 through the orifices 54 formed on the downstream flange 13 of the downstream turbine disk 5, this downstream flange 13 also participating in the outer delimitation of the annular space 88 and being assembled on the axis spacer 9 40.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (4)

  1. Vorrichtung zum Belüften eines Rotors einer HochdruckTurbine (100) einer Turbomaschine, wobei die Turbine (100) stromabwärts einer Verbrennungs-Kammer (2) angeordnet ist und umfasst: eine mit Schaufeln (4) versehene stromaufwärtige Turbinen-Scheibe (3) sowie eine mit Schaufeln (6) versehene stromabwärtige Turbinen-Scheibe (5), wobei die Vorrichtung umfasst: einen Kühl-Kreislauf, welcher mit Injektoren (36) versehen ist, welche stromaufwärts von der stromaufwärtigen Scheibe (3) angeordnet sind, und welche durch einen Durchfluss von Kühlluft (D) versorgt werden, welche(r) am Boden der Verbrennungs-Kammer (2) abgezweigt wird, wobei der Kühl-Kreislauf derart eingerichtet ist, dass der von den Injektoren (36) stammende Durchfluss von Kühlluft (D) Öffnungen (74) durchquert, welche in einem stromaufwärtigen Flansch (66) der stromaufwärtigen Scheibe (3) vorgesehen sind, was sein/ihr Fixieren auf einem stromaufwärtigen Flansch (78) der stromabwärtigen Scheibe (5) ermöglicht, damit dieser Durchfluss von Kühlluft (D) axial nach stromabwärts zwischen einer Innen-Bohrung (48) der stromaufwärtigen Scheibe (3) und dem stromaufwärtigen Flansch (78) der stromabwärtigen Scheibe (5) zirkuliert, was sein/ihr Fixieren auf einem stromabwärtigen Flansch (79) eines Hochdruck-Kompressors sowie das Zentrieren der stromaufwärtigen Scheibe (3) ermöglicht, wobei die Belüftungs-Vorrichtung ferner ein Einzel-Labyrinth (80) umfasst, welches mit einer der zwei Turbinen-Scheiben (3,5) verbunden ist,
    dadurch gekennzeichnet, das dieses Einzel-Labyrinth (80) zwischen den zwei Scheiben (3,5) zwischengeschaltet ist, derart, dass der Durchfluss von Kühlluft (D) sich aufteilt in einen ersten Fluss (F1), welcher zwischen einer stromabwärtigen Seite der stromaufwärtigen Scheibe (3) und einer stromaufwärtigen Seite des Einzel-Labyrinths (80) in Richtung der Schaufeln (4) zirkuliert, und in einen zweiten Fluss (F2), welcher zwischen einer stromaufwärtigen Seite der stromabwärtigen Scheibe (5) und einer stromabwärtigen Seite des Einzel-Labyrinths (80) in Richtung der Schaufeln (6) zirkuliert.
  2. Vorrichtung gemäß Anspruch 1, dadurch gekennzeichnet, dass die Injektoren (36) in eine Kavität (64) einmünden, welche teilweise von dem stromaufwärtigen Flansch (66) der stromaufwärtigen Turbinen-Scheibe (3) sowie durch eine stromaufwärtige Verbindung (32) und eine stromabwärtige Verbindung (34) begrenzt wird, wobei letztere mit einem zweiten stromaufwärtigen Flansch (72) der stromaufwärtigen Turbinen-Scheibe (3) zusammenwirkt.
  3. Vorrichtung gemäß Anspruch 1 oder Anspruch 2, dadurch gekennzeichnet, dass der stromaufwärtige Flansch (78) der stromabwärtigen Turbinen-Scheibe (5) eine Mehrzahl von Öffnungen (86) aufweist, welche dazu geeignet sind, von einem dritten Fluss (F3) des Kühl-Luft-Durchflusses (D) durchquert zu werden, wobei der dritte Fluss (F3) dazu geeignet ist, in einem ringförmigen Raum (88) axial nach stromabwärts zu zirkulieren, welcher zwischen einerseits dem stromaufwärtigen Flansch (78) der stromabwärtigen Scheibe (5) und einer Innen-Bohrung (50) dieser stromabwärtigen Scheibe (5), und andererseits einem rings einer Rotor-Welle (11) einer Niederdruck-Turbine angeordneten Distanzrohr/Abstandhalter (9) angeordnet ist.
  4. Vorrichtung gemäß einem der vorhergehenden Ansprüche, dadurch gekennzeichnet, dass das Einzel-Labyrinth (80) mit einem stromaufwärtigen zweiten Flansch (82) der stromabwärtigen Turbinen-Scheibe (5) verbunden ist, in welchem eine Mehrzahl von Öffnungen (84) eingearbeitet ist, welche die Zirkulation des zweiten Flusses (F2) des Kühl-Luft-Durchflusses (D) in Richtung der Schaufeln (6) ermöglichen.
EP04100404A 2003-02-06 2004-02-04 Vorrichtung für die Belüftung eines Rotor einer Hochdruckturbine Expired - Lifetime EP1445421B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0301391 2003-02-06
FR0301391A FR2851010B1 (fr) 2003-02-06 2003-02-06 Dispositif de ventilation d'un rotor de turbine a haute pression d'une turbomachine

Publications (2)

Publication Number Publication Date
EP1445421A1 EP1445421A1 (de) 2004-08-11
EP1445421B1 true EP1445421B1 (de) 2006-01-04

Family

ID=32606008

Family Applications (1)

Application Number Title Priority Date Filing Date
EP04100404A Expired - Lifetime EP1445421B1 (de) 2003-02-06 2004-02-04 Vorrichtung für die Belüftung eines Rotor einer Hochdruckturbine

Country Status (8)

Country Link
US (1) US6916151B2 (de)
EP (1) EP1445421B1 (de)
JP (1) JP4060279B2 (de)
CA (1) CA2456589C (de)
DE (1) DE602004000301T2 (de)
ES (1) ES2255697T3 (de)
FR (1) FR2851010B1 (de)
RU (1) RU2330976C2 (de)

Families Citing this family (23)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2892148B1 (fr) * 2005-10-19 2011-07-22 Snecma Fourreau d'arbre de turboreacteur et turboreacteur comportant ce fourreau
US8668437B1 (en) * 2006-09-22 2014-03-11 Siemens Energy, Inc. Turbine engine cooling fluid feed system
US8562285B2 (en) * 2007-07-02 2013-10-22 United Technologies Corporation Angled on-board injector
FR2937371B1 (fr) * 2008-10-20 2010-12-10 Snecma Ventilation d'une turbine haute-pression dans une turbomachine
FR2946687B1 (fr) * 2009-06-10 2011-07-01 Snecma Turbomachine comprenant des moyens ameliores de reglage du debit d'un flux d'air de refroidissement preleve en sortie de compresseur haute pression
US8371127B2 (en) * 2009-10-01 2013-02-12 Pratt & Whitney Canada Corp. Cooling air system for mid turbine frame
FR2960260B1 (fr) * 2010-05-21 2014-05-09 Snecma Turbomachine comprenant un circuit de ventilation de turbine basse pression ameliore
JP5494457B2 (ja) * 2010-12-13 2014-05-14 トヨタ自動車株式会社 ガスタービンエンジン
US20120308360A1 (en) * 2011-05-31 2012-12-06 General Electric Company Overlap seal for turbine nozzle assembly
US9279341B2 (en) 2011-09-22 2016-03-08 Pratt & Whitney Canada Corp. Air system architecture for a mid-turbine frame module
US9091173B2 (en) * 2012-05-31 2015-07-28 United Technologies Corporation Turbine coolant supply system
US20130327061A1 (en) * 2012-06-06 2013-12-12 General Electric Company Turbomachine bucket assembly and method of cooling a turbomachine bucket assembly
US10167723B2 (en) * 2014-06-06 2019-01-01 United Technologies Corporation Thermally isolated turbine section for a gas turbine engine
US9915204B2 (en) * 2014-06-19 2018-03-13 United Technologies Corporation Systems and methods for distributing cooling air in gas turbine engines
CN104675447A (zh) * 2015-01-30 2015-06-03 北京华清燃气轮机与煤气化联合循环工程技术有限公司 一种燃气轮机涡轮冷却气路
US10634055B2 (en) * 2015-02-05 2020-04-28 United Technologies Corporation Gas turbine engine having section with thermally isolated area
US9920652B2 (en) 2015-02-09 2018-03-20 United Technologies Corporation Gas turbine engine having section with thermally isolated area
JP2018514685A (ja) * 2015-04-06 2018-06-07 シーメンス エナジー インコーポレイテッド タービン翼の二圧力冷却
US10718220B2 (en) * 2015-10-26 2020-07-21 Rolls-Royce Corporation System and method to retain a turbine cover plate with a spanner nut
US10030519B2 (en) * 2015-10-26 2018-07-24 Rolls-Royce Corporation System and method to retain a turbine cover plate between nested turbines with a tie bolt and spanner nut
US10273812B2 (en) 2015-12-18 2019-04-30 Pratt & Whitney Canada Corp. Turbine rotor coolant supply system
US11421597B2 (en) 2019-10-18 2022-08-23 Pratt & Whitney Canada Corp. Tangential on-board injector (TOBI) assembly
CN111946464B (zh) * 2020-07-21 2021-09-07 中国科学院工程热物理研究所 一种用于高压涡轮盘后轴承腔的导流阻挡密封结构

Family Cites Families (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3043561A (en) * 1958-12-29 1962-07-10 Gen Electric Turbine rotor ventilation system
GB2081392B (en) * 1980-08-06 1983-09-21 Rolls Royce Turbomachine seal
US4462204A (en) * 1982-07-23 1984-07-31 General Electric Company Gas turbine engine cooling airflow modulator
US4882902A (en) * 1986-04-30 1989-11-28 General Electric Company Turbine cooling air transferring apparatus
GB2189845B (en) * 1986-04-30 1991-01-23 Gen Electric Turbine cooling air transferring apparatus
FR2712029B1 (fr) * 1993-11-03 1995-12-08 Snecma Turbomachine pourvue d'un moyen de réchauffage des disques de turbines aux montées en régime.
US5555721A (en) * 1994-09-28 1996-09-17 General Electric Company Gas turbine engine cooling supply circuit
DE19854907A1 (de) * 1998-11-27 2000-05-31 Rolls Royce Deutschland Kühlluftführung an einer Axialturbine
US6468032B2 (en) * 2000-12-18 2002-10-22 Pratt & Whitney Canada Corp. Further cooling of pre-swirl flow entering cooled rotor aerofoils
US6540477B2 (en) * 2001-05-21 2003-04-01 General Electric Company Turbine cooling circuit
US6735956B2 (en) * 2001-10-26 2004-05-18 Pratt & Whitney Canada Corp. High pressure turbine blade cooling scoop

Also Published As

Publication number Publication date
RU2004103479A (ru) 2005-07-10
CA2456589A1 (en) 2004-08-06
DE602004000301D1 (de) 2006-03-30
CA2456589C (en) 2012-04-24
JP2004239260A (ja) 2004-08-26
US6916151B2 (en) 2005-07-12
RU2330976C2 (ru) 2008-08-10
DE602004000301T2 (de) 2006-08-31
ES2255697T3 (es) 2006-07-01
JP4060279B2 (ja) 2008-03-12
FR2851010A1 (fr) 2004-08-13
EP1445421A1 (de) 2004-08-11
US20040219008A1 (en) 2004-11-04
FR2851010B1 (fr) 2005-04-15

Similar Documents

Publication Publication Date Title
EP1445421B1 (de) Vorrichtung für die Belüftung eines Rotor einer Hochdruckturbine
EP2440746B1 (de) Turbinenmotor mit verbesserter vorrichtung zur einstellung der strömungsgeschwindigkeit eines am ausgang eines hochdruckverdichters abgezapften kühlluftstroms
CA2758175C (fr) Moteur a turbine a gaz a double corps pourvu d ' un palier inter-arbres
EP2337929B1 (de) Lüftung einer hochdruckturbine in einer turbomaschine
CA2564491C (fr) Dispositif de ventilation de disques de turbine dans un moteur a turbine a gaz
CA2475404C (fr) Echangeur sur circuit de ventilation de turbine
EP1316675B1 (de) Stator für eine Turbomaschine
CA2641963C (fr) Controle du jeu en sommet d'aubes dans une turbine haute-pression de turbomachine
CA2430143C (fr) Refroidissement du flasque amont d'une turbine a haute pression par un systeme a double injecteurs fond de chambre
BE1023354B1 (fr) Bec de separation degivrant de compresseur de turbomachine axiale
EP1552111A1 (de) Gasturbinenlüftungsschaltungsanordnung
FR2598179A1 (fr) Dispositif de transfert d'air de refroidissement pour une turbine
EP2078822B1 (de) Gasturbinenmotor mit Ventil für die Verbindung von zwei Behältern
WO2021191523A1 (fr) Turbomachine avec dispositif de refroidissement et de pressurisation d'une turbine
FR2999249A1 (fr) Compresseur pour turbomachine dote de moyens de refroidissement d'un joint tournant assurant l'etancheite entre un redresseur et un rotor
WO2023047055A1 (fr) Carter d'injection d'air de refroidissement pour turbine de turbomachine
FR2881472A1 (fr) Circuit de ventilation d'un rotor de turbine haute pression dans un moteur a turbine a gaz
FR2960020A1 (fr) Dispositif et procede de refroidissement de disques de rotor d'une turbomachine
FR2557634A2 (fr) Dispositif d'etancheite d'aubages mobiles de turbomachine

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HU IE IT LI LU MC NL PT RO SE SI SK TR

AX Request for extension of the european patent

Extension state: AL LT LV MK

17P Request for examination filed

Effective date: 20050131

AKX Designation fees paid

Designated state(s): DE ES FR GB IT SE

17Q First examination report despatched

Effective date: 20050527

GRAP Despatch of communication of intention to grant a patent

Free format text: ORIGINAL CODE: EPIDOSNIGR1

RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: SNECMA

GRAS Grant fee paid

Free format text: ORIGINAL CODE: EPIDOSNIGR3

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): DE ES FR GB IT SE

REG Reference to a national code

Ref country code: GB

Ref legal event code: FG4D

Free format text: NOT ENGLISH

GBT Gb: translation of ep patent filed (gb section 77(6)(a)/1977)

Effective date: 20060213

REF Corresponds to:

Ref document number: 602004000301

Country of ref document: DE

Date of ref document: 20060330

Kind code of ref document: P

REG Reference to a national code

Ref country code: SE

Ref legal event code: TRGR

REG Reference to a national code

Ref country code: ES

Ref legal event code: FG2A

Ref document number: 2255697

Country of ref document: ES

Kind code of ref document: T3

PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

26N No opposition filed

Effective date: 20061005

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: ES

Payment date: 20130207

Year of fee payment: 10

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 13

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: IT

Payment date: 20160127

Year of fee payment: 13

Ref country code: DE

Payment date: 20160121

Year of fee payment: 13

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: SE

Payment date: 20160126

Year of fee payment: 13

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: ES

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20140205

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 14

REG Reference to a national code

Ref country code: DE

Ref legal event code: R119

Ref document number: 602004000301

Country of ref document: DE

REG Reference to a national code

Ref country code: SE

Ref legal event code: EUG

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: SE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20170205

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 15

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: DE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20170901

REG Reference to a national code

Ref country code: FR

Ref legal event code: CD

Owner name: SAFRAN AIRCRAFT ENGINES

Effective date: 20170719

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: IT

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20170204

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: FR

Payment date: 20230119

Year of fee payment: 20

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: GB

Payment date: 20230121

Year of fee payment: 20

REG Reference to a national code

Ref country code: GB

Ref legal event code: PE20

Expiry date: 20240203

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: GB

Free format text: LAPSE BECAUSE OF EXPIRATION OF PROTECTION

Effective date: 20240203