EP1653047B1 - Aube de rotor d'une turbine à gaz - Google Patents

Aube de rotor d'une turbine à gaz Download PDF

Info

Publication number
EP1653047B1
EP1653047B1 EP20050292209 EP05292209A EP1653047B1 EP 1653047 B1 EP1653047 B1 EP 1653047B1 EP 20050292209 EP20050292209 EP 20050292209 EP 05292209 A EP05292209 A EP 05292209A EP 1653047 B1 EP1653047 B1 EP 1653047B1
Authority
EP
European Patent Office
Prior art keywords
blade
cavity
platform
stiffener
turbine
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
EP20050292209
Other languages
German (de)
English (en)
French (fr)
Other versions
EP1653047A2 (fr
EP1653047A3 (fr
Inventor
Jacques Boury
Maurice 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 EP1653047A2 publication Critical patent/EP1653047A2/fr
Publication of EP1653047A3 publication Critical patent/EP1653047A3/fr
Application granted granted Critical
Publication of EP1653047B1 publication Critical patent/EP1653047B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • 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
    • 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
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms

Definitions

  • the present invention relates to a rotor blade of a gas turbine, in particular a high-pressure turbojet turbine.
  • a rotor blade of a gas turbine comprises a blade formed with an extrados or convex outer surface and with a concave inner surface or inner surface, connected at their upstream ends by a leading edge and at their downstream ends. by a trailing edge of the gases.
  • the blade is connected by a platform to a blade root dovetail, fir or the like, intended to be inserted into a corresponding cavity of a rotor disc of the gas turbine.
  • At least one reinforcing web called “stiffener" is formed at the downstream end of the platform on the side opposite to the blade and extends transversely while being connected to the blade root.
  • the blade also comprises cooling means by circulating a fluid such as air in ducts formed of foundry inside the blade and the blade root.
  • the cooling air exits in particular through discharge slots which open downstream along the trailing edge and which are oriented substantially perpendicular to the longitudinal axis of the blade and parallel to the platform.
  • connection area of the trailing edge to the platform is between a cooling air discharge slot and the stiffener, the radially inner portion is cooled by contact with the cooling air.
  • This connection zone in contact with the hot gases passing through the turbine, is subjected to intense thermal stresses which cause the formation of cracks likely to destroy the blade and also the turbine.
  • the invention aims in particular to provide a simple, economical and effective solution to this problem.
  • connection zone between the trailing edge and the platform is cooled by limiting the thermal gradient between this connection zone and the stiffener.
  • the cooling cavity formed in the stiffener substantially to the right of the trailing edge to cool the material between the cavity and the connection of the trailing edge to the platform. This results in a significant reduction in the thermal gradient between this connection and the stiffener and in a concomitant reduction in the risk of formation of cracks at the connection of the trailing edge to the platform.
  • the outlet or openings of the cavity are advantageously substantially parallel to the trailing edge. They allow an output of the cooling fluid circulating in the stiffener cavity without disturbing the flow of gases leaving the blade.
  • the stiffener cavity is realizable from foundry with ducts coolant circulation and the cavity outlet ports are also obtained by casting when they have a diameter greater than or equal to about 0.6 millimeters, or are made by laser drilling or by electro-erosion when they have a smaller diameter .
  • the stiffener can be given a thickness slightly greater than that normally expected, the increase in mass due to this increase in thickness being compensated for by the formation of the cavity.
  • the invention also proposes a turbojet turbine, characterized in that it comprises a plurality of blades of the above-mentioned type, the stiffeners of which are formed with cooling cavities substantially in line with the trailing edges of the vanes.
  • the invention also relates to a turbojet, characterized in that it comprises a turbine as described above.
  • FIG. 1 is represented a blade 10 of a high-pressure stage of a gas turbine, in particular of a turbojet engine.
  • This blade 10 comprises a blade formed with an extrados 12 or convex outer surface and with a lower surface 14 or concave inner surface which are connected at their upstream ends by a leading edge 16 and at their downstream ends by a trailing edge 18 of gas flowing into the turbine.
  • the blade is connected by a substantially rectangular transverse platform 20 to a blade root 22 by means of which the blade 10 is mounted on a disc (not shown) of the rotor of the gas turbine, by fitting of this foot 22 into a correspondingly shaped cavity of the periphery of the rotor disc.
  • this male / female fitting which is fir type in the example shown, the blade 10 is retained radially on the rotor disc.
  • Other means are provided for axially locking the root 22 of the blade 10 in the cavity of the disc.
  • Each rotor disk comprises a plurality of vanes 10 regularly distributed on its outer periphery.
  • the platform 20 is also connected to the blade root 22 by reinforcing webs called stiffeners 24, 26 which extend on the opposite side to the blade at the upstream and downstream ends of the platform 20, respectively, substantially perpendicular to the platform 20 and transversely or circumferentially with respect to the axis of rotation when the blade 10 is mounted on a rotor disc.
  • the downstream stiffener 26 extends under the junction between the trailing edge 18 and the platform 20 and is connected to the blade root 22. Its lateral edge 28 substantially perpendicular to the platform 10 connects its radially inner edge 30 to a lateral edge platform 20 at the junction between the trailing edge 18 and the platform 20.
  • the upstream and downstream stiffeners 24, 26 stiffen the platform 20 and prevent it from bending outwards about an axis parallel to the axis of rotation, and delimit between them a housing of a sealing jacket (not shown) which is arranged under the platform 20 and which extends between this blade 10 and an adjacent blade of the rotor disc.
  • sealing sleeves prevent the passage of gas or air from the inner part of the turbine radially outwards between the platforms 20 of the adjacent blades, and conversely prevent the passage of gas or air from the outside towards the inner part of the turbine between the platforms 20 of the adjacent blades.
  • the air of the inner part engages in orifices 32 of the end face of the blade root 22 and circulates in supply ducts 34 formed in the blade root 22 and extending into the airfoil. dawn 10, as indicated by dotted lines in figure 2 these ducts being substantially parallel to the longitudinal axis 44 of the blade 10 and serving for cooling thereof.
  • the circulation of air in the supply ducts is schematically represented by dashed arrows.
  • the channel 34 located near the trailing edge 18 of the blade 10 feeds air exhaust slots 46, shown in FIG. figure 1 and delimited in figure 2 by dotted lines, formed on a portion of the intrados 14 near the trailing edge 18 and oriented substantially perpendicular to the longitudinal axis 44 of the blade 10 and parallel to the platform 20.
  • the cooling air exiting through the slots 46 of the trailing edge 18 can not cool the connection 48 between the trailing edge 18 and the platform 20, which is in contact with the hot gases and is subjected to thermal stresses important.
  • the invention provides for reducing these constraints by reducing the vertical thermal gradient between the downstream stiffener 26 and the connection 48 of the trailing edge 18 to the platform 20.
  • a cavity 50 is formed in the stiffener 26, substantially in line with the trailing edge 18, and communicates with a duct 34 for supplying cooling air and with means for outputting the cooling air.
  • the cavity 50 has a substantially parallelepipedal shape with an inner edge 52 near the inner edge 30 of the stiffener 26 and substantially parallel thereto, a lateral edge 54 near the lateral edge 28 of the stiffener 26 and substantially parallel to the latter. ci, and an outer edge 56 substantially adjacent to the platform 20.
  • the cavity 50 is connected directly to the conduit 34 for supplying the discharge slots 46 with cooling air.
  • the cavity 50 is connected to the outside by one or more orifices 58 opening downstream under the platform, which make it possible to ensure a continuous flow of air inside the cavity 50 and to cool the material located between this cavity 50 and the connection 48 of the trailing edge 18 to the platform 20.
  • the circulation of the air in the cavity 50 and its outlet through the orifices 58 carries out a transfer and a heat discharge from the material between the cavity 50 and the connection 48 of the trailing edge 18, and cools this connection 48 by conduction.
  • These orifices 58 may be of any shape and size. They can be formed on the downstream face of the stiffener 26.
  • the cavity 50 has a length of about 5 to 6 millimeters in transverse or circumferential dimension, a height of about 3 millimeters along the axis 44 of dawn, and a thickness of 1 millimeter or less, for example about 0.8 millimeters, along the axis of rotation.
  • This cavity 50 is advantageously made of foundry. In order not to weaken the downstream stiffener 26 of the blade 10, the thickness thereof can be increased, the increase in mass due to this increase in thickness being compensated for by the formation of the cavity 50.
  • the orifices 58 are made of foundry, by laser drilling or electroerosion, the laser drilling techniques and EDM being substituted for the foundry for the production of orifices with a diameter less than about 0.6 millimeter.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP20050292209 2004-10-27 2005-10-20 Aube de rotor d'une turbine à gaz Active EP1653047B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0411436A FR2877034B1 (fr) 2004-10-27 2004-10-27 Aube de rotor d'une turbine a gaz

Publications (3)

Publication Number Publication Date
EP1653047A2 EP1653047A2 (fr) 2006-05-03
EP1653047A3 EP1653047A3 (fr) 2011-09-07
EP1653047B1 true EP1653047B1 (fr) 2015-04-29

Family

ID=34952822

Family Applications (1)

Application Number Title Priority Date Filing Date
EP20050292209 Active EP1653047B1 (fr) 2004-10-27 2005-10-20 Aube de rotor d'une turbine à gaz

Country Status (4)

Country Link
US (1) US7497661B2 (enrdf_load_stackoverflow)
EP (1) EP1653047B1 (enrdf_load_stackoverflow)
JP (1) JP4663479B2 (enrdf_load_stackoverflow)
FR (1) FR2877034B1 (enrdf_load_stackoverflow)

Families Citing this family (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8147197B2 (en) * 2009-03-10 2012-04-03 Honeywell International, Inc. Turbine blade platform
US8133024B1 (en) 2009-06-23 2012-03-13 Florida Turbine Technologies, Inc. Turbine blade with root corner cooling
US8550783B2 (en) 2011-04-01 2013-10-08 Alstom Technology Ltd. Turbine blade platform undercut
JP2011241836A (ja) * 2011-08-02 2011-12-01 Mitsubishi Heavy Ind Ltd ガスタービン動翼のプラットフォーム冷却構造
CN102418562B (zh) * 2011-08-15 2014-04-02 清华大学 一种纤维缠绕的预应力涡轮转子
US8840370B2 (en) 2011-11-04 2014-09-23 General Electric Company Bucket assembly for turbine system
US8870525B2 (en) 2011-11-04 2014-10-28 General Electric Company Bucket assembly for turbine system
US8845289B2 (en) 2011-11-04 2014-09-30 General Electric Company Bucket assembly for turbine system
US8974182B2 (en) * 2012-03-01 2015-03-10 General Electric Company Turbine bucket with a core cavity having a contoured turn
RU2014149236A (ru) 2012-05-08 2016-06-27 Сименс Акциенгезелльшафт Лопатка ротора турбины и осевой участок ротора для газовой турбины
CN105855468A (zh) * 2016-04-13 2016-08-17 东方电气集团东方汽轮机有限公司 陶瓷型壳制备方法及制备透平叶片的陶瓷型壳的制造方法
US11021961B2 (en) 2018-12-05 2021-06-01 General Electric Company Rotor assembly thermal attenuation structure and system
CN112459849B (zh) * 2020-10-27 2022-08-30 哈尔滨广瀚燃气轮机有限公司 一种用于燃气轮机涡轮叶片的冷却结构

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS5979006A (ja) * 1982-10-27 1984-05-08 Hitachi Ltd ガスタ−ビン空冷翼

Family Cites Families (19)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4244676A (en) 1979-06-01 1981-01-13 General Electric Company Cooling system for a gas turbine using a cylindrical insert having V-shaped notch weirs
WO1996006266A1 (en) * 1994-08-24 1996-02-29 Westinghouse Electric Corporation Gas turbine blade with cooled platform
KR100364183B1 (ko) 1994-10-31 2003-02-19 웨스팅하우스 일렉트릭 코포레이션 냉각된플랫폼을구비한가스터빈블레이드
JP3758792B2 (ja) * 1997-02-25 2006-03-22 三菱重工業株式会社 ガスタービン動翼のプラットフォーム冷却機構
JP3316418B2 (ja) * 1997-06-12 2002-08-19 三菱重工業株式会社 ガスタービン冷却動翼
CA2262064C (en) * 1998-02-23 2002-09-03 Mitsubishi Heavy Industries, Ltd. Gas turbine moving blade platform
JP3426952B2 (ja) * 1998-03-03 2003-07-14 三菱重工業株式会社 ガスタービン動翼のプラットフォーム
US6190130B1 (en) * 1998-03-03 2001-02-20 Mitsubishi Heavy Industries, Ltd. Gas turbine moving blade platform
CA2231988C (en) * 1998-03-12 2002-05-28 Mitsubishi Heavy Industries, Ltd. Gas turbine blade
US6210111B1 (en) * 1998-12-21 2001-04-03 United Technologies Corporation Turbine blade with platform cooling
KR100694370B1 (ko) * 1999-05-14 2007-03-12 제너럴 일렉트릭 캄파니 터빈 노즐의 내측 및 외측 밴드에서 온도 부정합을 제어하는 장치 및 내측 또는 외측 밴드의 벽과 커버 사이의 온도 차이를 감소시키는 방법
US6390774B1 (en) * 2000-02-02 2002-05-21 General Electric Company Gas turbine bucket cooling circuit and related process
FR2835015B1 (fr) * 2002-01-23 2005-02-18 Snecma Moteurs Aube mobile de turbine haute pression munie d'un bord de fuite au comportement thermique ameliore
GB2395987B (en) * 2002-12-02 2005-12-21 Alstom Turbine blade with cooling bores
JP3776897B2 (ja) * 2003-07-31 2006-05-17 三菱重工業株式会社 ガスタービン動翼のプラットフォーム冷却機構
US6923616B2 (en) * 2003-09-02 2005-08-02 General Electric Company Methods and apparatus for cooling gas turbine engine rotor assemblies
US6945749B2 (en) * 2003-09-12 2005-09-20 Siemens Westinghouse Power Corporation Turbine blade platform cooling system
US7097417B2 (en) * 2004-02-09 2006-08-29 Siemens Westinghouse Power Corporation Cooling system for an airfoil vane
GB0405679D0 (en) * 2004-03-13 2004-04-21 Rolls Royce Plc A mounting arrangement for turbine blades

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS5979006A (ja) * 1982-10-27 1984-05-08 Hitachi Ltd ガスタ−ビン空冷翼

Also Published As

Publication number Publication date
JP4663479B2 (ja) 2011-04-06
EP1653047A2 (fr) 2006-05-03
US7497661B2 (en) 2009-03-03
JP2006125402A (ja) 2006-05-18
FR2877034A1 (fr) 2006-04-28
FR2877034B1 (fr) 2009-04-03
US20060088416A1 (en) 2006-04-27
EP1653047A3 (fr) 2011-09-07

Similar Documents

Publication Publication Date Title
EP1653047B1 (fr) Aube de rotor d'une turbine à gaz
EP1288439B1 (fr) Circuit de refroidissement pour aube de turbine à gaz
EP0666406B1 (fr) Aube fixe ou mobile refroidie de turbine
EP1736636B1 (fr) Aube creuse de turbomachine
CA2550442C (fr) Circuits de refroidissement pour aube mobile de turbomachine
CA2652679C (fr) Aubes pour roue a aubes de turbomachine avec rainure pour le refroidissement
EP3073053B1 (fr) Aube mobile de turbine a conception amelioree pour turbomachine d'aeronef
EP2598721B1 (fr) Roue de turbine ou de compresseur de turbomachine
FR2981979A1 (fr) Roue de turbine pour une turbomachine
FR2943091A1 (fr) Aube de turbine avec un trou de depoussierage en base de pale
CA2619299A1 (fr) Disque de rotor de soufflante de turbomachine
WO2015162389A1 (fr) Aube pour turbine de turbomachine comprenant un circuit de refroidissement à homogénéité améliorée
FR2974387A1 (fr) Roue de turbine pour une turbomachine
CA2456705C (fr) Plate-forme annulaire de distributeur d'une turbine basse-pression de turbomachine
EP1630351B1 (fr) Aube de compresseur ou de turbine à gaz
EP1630350B1 (fr) Aube de rotor d'un compresseur ou d'une turbine à gaz
FR3044038B1 (fr) Aube equipee d'un systeme de refroidissement, distributeur et turbomachine associes
WO2020193913A1 (fr) Aube de turbomachine equipee d'un circuit de refroidissement optimise
EP3712378A1 (fr) Aube de turbomachine, comportant des deflecteurs dans une cavite interne de refroidissement
WO2020120862A1 (fr) Aube de turbomachine à refroidissement amélioré
WO2020193912A1 (fr) Aube de turbomachine equipee d'un circuit de refroidissement et procede de fabrication a cire perdue d'une telle aube

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: A2

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

AX Request for extension of the european patent

Extension state: AL BA HR MK YU

PUAL Search report despatched

Free format text: ORIGINAL CODE: 0009013

AK Designated contracting states

Kind code of ref document: A3

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

AX Request for extension of the european patent

Extension state: AL BA HR MK YU

RIC1 Information provided on ipc code assigned before grant

Ipc: F01D 5/08 20060101ALI20110801BHEP

Ipc: F01D 5/18 20060101AFI20110801BHEP

17P Request for examination filed

Effective date: 20120209

AKX Designation fees paid

Designated state(s): DE FR GB

17Q First examination report despatched

Effective date: 20120613

GRAP Despatch of communication of intention to grant a patent

Free format text: ORIGINAL CODE: EPIDOSNIGR1

INTG Intention to grant announced

Effective date: 20141117

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 FR GB

REG Reference to a national code

Ref country code: GB

Ref legal event code: FG4D

Free format text: NOT ENGLISH

REG Reference to a national code

Ref country code: DE

Ref legal event code: R096

Ref document number: 602005046435

Country of ref document: DE

Effective date: 20150603

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 11

REG Reference to a national code

Ref country code: DE

Ref legal event code: R097

Ref document number: 602005046435

Country of ref document: DE

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: 20160201

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 12

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 13

REG Reference to a national code

Ref country code: FR

Ref legal event code: CD

Owner name: SAFRAN AIRCRAFT ENGINES, FR

Effective date: 20170719

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 14

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

Ref country code: GB

Payment date: 20240919

Year of fee payment: 20

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

Ref country code: FR

Payment date: 20240919

Year of fee payment: 20

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

Ref country code: DE

Payment date: 20240919

Year of fee payment: 20