EP1950381B1 - Disque de rotor de soufflante de turbomachine - Google Patents

Disque de rotor de soufflante de turbomachine Download PDF

Info

Publication number
EP1950381B1
EP1950381B1 EP07291632.3A EP07291632A EP1950381B1 EP 1950381 B1 EP1950381 B1 EP 1950381B1 EP 07291632 A EP07291632 A EP 07291632A EP 1950381 B1 EP1950381 B1 EP 1950381B1
Authority
EP
European Patent Office
Prior art keywords
disk
blade
cavities
grooves
turbomachine
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
EP07291632.3A
Other languages
German (de)
English (en)
French (fr)
Other versions
EP1950381A1 (fr
Inventor
Son Le Hong
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 EP1950381A1 publication Critical patent/EP1950381A1/fr
Application granted granted Critical
Publication of EP1950381B1 publication Critical patent/EP1950381B1/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
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
    • 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
    • F01D21/00—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
    • F01D21/04—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position
    • F01D21/045—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position special arrangements in stators or in rotors dealing with breaking-off of part of rotor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F04D29/322—Blade mountings

Definitions

  • the present invention relates to a fan rotor disk for a turbomachine, such in particular as an airplane turbojet engine.
  • a fan rotor disc comprises a plurality of vanes mounted at its periphery and separated from each other by platforms fixed to flanges of the disc.
  • Each blade is formed of a blade connected to a blade root by means of a stilt.
  • the blade roots are engaged in substantially axial grooves, formed at the periphery of the disk and are held there radially by cooperation of shapes, the blade roots being for example dovetail cross section or the like.
  • breaking the connection of a blade with the disk can cause the destruction of neighboring blades and adjoining platforms. Indeed, in case of loss of fan blade, it is supported on the next blade, and the resulting force applied to the blade is reflected in particular by an axial stress directed from the downstream to the upstream due to the angular setting of the blade relative to the groove, which tends to tilt the blade upstream and generate a strong constraint at the rear link between the blade root and the disc. A breakage of the blade root or tooth of the disk can thus occur, leading to a chain reaction that can destroy all the blades of the fan and the platforms and strongly damaging the turbomachine.
  • the blade root which is engaged in the groove is connected downstream to a hook.
  • Notches formed radially on either side of each hook cooperate with an annular flange to ensure the axial retention of the blades when they are positioned in the grooves of the disc.
  • this method of attachment generates a strong constraint at the hook stitch connection area and at the connection of the notch with the hook. As before, this constraint can cause a break, at the level of the hook of the blade or at the level of the disk, and cause a chain destruction of the blades and platforms.
  • an axial groove about 10 mm long, opening on the notch, is machined on each side of the blade root, to limit the stress applied at the connection area of the stilt with the hook and at the connection area between the notch and the hook, by directing the forces upstream of the machining.
  • This groove if it makes it possible to limit the forces at the hook, nevertheless has the disadvantage of generating a peak of stress at its upstream end, which causes a large wear of the blade root and the disk and thus limits their duration. life.
  • Several solutions have been considered to limit the wear of these parts and consisted in forming a relief at the upstream end of the machining, or to place a foil between the blade and the disk. However, these means do not solve satisfactorily the wear problem while limiting the stress applied to the hook of the blade and transmitted to the platforms.
  • the document GB-A-2100808 describes a fan blade comprising a dovetail foot engaged in a groove of a disk.
  • the foot of the blade has at its downstream end reinforcing means to stiffen it.
  • the shock energy is transmitted to the disk and the disk areas located radially outside the groove will deform elastically or plastically.
  • the document GB-A-23800770 describes a turbine blade engaged in a groove of a disk which comprises cavities located radially outside the grooves.
  • the cavities are open laterally and open into the grooves.
  • the shape of the cavities is optimized to avoid the formation of stress peaks due to the flight forces of the blade roots on the groove of the disk. These cavities are also able to deform in case of dawn loss.
  • the document US-B1-6634863 describes a fan disk having at its periphery a plurality of grooves in which are engaged axially the blade roots.
  • Each blade root comprises surfaces in contact with corresponding surfaces of the groove of the disk.
  • the groove includes recesses to reduce stresses in flight of the blade root on the groove at the contact surfaces.
  • the invention aims in particular to provide a simple, economical and effective solution to these various problems.
  • a turbomachine comprising a fan rotor disk for a turbomachine, comprising on the periphery substantially axial grooves intended for the mounting and radial retention of blade roots, the disk comprising deformable zones in the event of loss of blade formed by cavities located at the downstream end of the grooves, characterized in that the cavities are formed in flanges intended for the attachment of inter-blade platforms, these flanges extending in the extension of the side walls of the grooves.
  • the blades of the rotor disc according to the invention no longer require axial machining allowing the deviation of forces. This eliminates the phenomena of wear of the disk and the blade due to this machining while limiting the stresses applied to the hooks and transmitted to the platforms, thanks to the cavities formed in the clamping straps of international platforms. blades.
  • the cavities are formed by machining.
  • the cavities are oriented axially and are of closed bottom tubular shape.
  • the cavities are formed by drilling or milling.
  • the cavities are open laterally and open into the grooves.
  • the invention also relates to a turbomachine, such as an aircraft turbojet engine, characterized in that it comprises a fan rotor disc of the type described above.
  • FIG 1 representing a fan disc 10 carrying a blade 12 and the figure 2 which represents the radially inner downstream part of a blade according to the prior art.
  • a blade is formed of a blade 14 connected to a blade root 20 via a stagger 18.
  • the disk 10 has a plurality of substantially axial grooves 22 regularly distributed at its outer periphery and in which the blades are engaged. blades 12. Platforms (not shown) are arranged between the blades and serve to direct the flow of air at the inlet of the turbomachine.
  • the dovetail-shaped dovetail 20 or the like cooperates with the groove 22 to provide radial retention of the vane (12) on the rotor disc 10.
  • a hook 24 comprising a radial notch 26 on each of its lateral faces. These notches cooperate with an annular flange 28 to axially block the foot 20 of the blade 12 in the groove 22 of the disk 10.
  • the hook / hook 30 and notch / hook connection zones 32 are strongly stressed.
  • the radial contact of the blade disconnected from the disk with the neighboring blade results in the attachment of the blade in a groove by additional stress in the stitch / hook connection zones. and notch / hook 32. Therefore, the stress applied to the rear of the blade weakens the hook 24, which may cause its break.
  • Such a constraint can also damage the disk and therefore the inter-blade platforms that are attached thereto.
  • the rupture of the connection with the disk of a second blade can cause a chain reaction leading to the complete destruction of the fan blades and adjoining platforms, leading to significant damage to the turbomachine. It is therefore imperative to maintain the blades in position in their grooves as well as the platforms on the flanges of attachment of the disk in the event of loss of blades.
  • an axial machining 38 is performed on each side of the hook 24, and opens on the notch 26.
  • the axial machining 38 allows to offset the forces, represented by dotted arrows, beyond the machining which reduces the constraints applied to the hook, the forces in the absence of machining being represented in solid arrows.
  • the constraints applied to the hook are thus limited and the dawn has a better hold.
  • this type of solution is not satisfactory since a high stress is generated at the upstream end of the machining 38, which causes a significant wear of the blade root and the disc.
  • the invention proposes, to eliminate this phenomenon of wear while limiting the stress applied to the connection of the blade with the disk and transmitted to the platforms, to form deformable zones 34 in the disk 10 located radially to the outside the grooves 22, at the hooks of the blade roots.
  • deformable zones 34 are formed by cavities 34 made in hooking flanges 36 of inter-blade platforms (not shown), and are fixed on flanges 36 extending substantially in the extension of the side walls of the grooves 22 ( figure 3 to 5 )
  • the cavities 34 are open laterally and open into the grooves.
  • the cavity has for example a diameter of the order of 6 to 9 mm, the thickness of the cavity wall is between 0 and 3 mm, the depth being about 20 mm.
  • cavities can be made by simple and fast machining techniques such as drilling or milling.
  • cavities 34 in the clamping flanges 36 of the inter-blade platforms allows the plastic deformation of these cavities in case of dawn loss.
  • the forces at the outlet of the blade reach are directed towards the cavities 34.
  • the stress applied to the rear hook is lower, which prevents the hook from breaking and allows the blade to remain in position in its groove and the plates. adjoining forms to remain fixed on the flanges 36 of the disc 10 until the shutdown of the turbomachine.
  • the service life is no longer limited by wear phenomena due to axial machining in the blade root 20, the latter being no longer necessary.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP07291632.3A 2007-01-18 2007-12-27 Disque de rotor de soufflante de turbomachine Active EP1950381B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0700326A FR2911632B1 (fr) 2007-01-18 2007-01-18 Disque de rotor de soufflante de turbomachine

Publications (2)

Publication Number Publication Date
EP1950381A1 EP1950381A1 (fr) 2008-07-30
EP1950381B1 true EP1950381B1 (fr) 2016-03-02

Family

ID=38421439

Family Applications (1)

Application Number Title Priority Date Filing Date
EP07291632.3A Active EP1950381B1 (fr) 2007-01-18 2007-12-27 Disque de rotor de soufflante de turbomachine

Country Status (6)

Country Link
US (1) US8246309B2 (enExample)
EP (1) EP1950381B1 (enExample)
JP (1) JP5283388B2 (enExample)
CA (1) CA2619299C (enExample)
FR (1) FR2911632B1 (enExample)
RU (1) RU2454572C2 (enExample)

Families Citing this family (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
TWM334886U (en) * 2007-12-12 2008-06-21 Taiwei Fan Technology Co Ltd Combination type miniature axial-flow fan
DE102009007468A1 (de) * 2009-02-04 2010-08-19 Mtu Aero Engines Gmbh Integral beschaufelte Rotorscheibe für eine Turbine
US8485784B2 (en) * 2009-07-14 2013-07-16 General Electric Company Turbine bucket lockwire rotation prevention
EP2299056A1 (de) * 2009-09-02 2011-03-23 Siemens Aktiengesellschaft Kühlung eines Gasturbinenbauteils ausgebildet als Rotorscheibe oder Turbinenschaufel
FR2955904B1 (fr) * 2010-02-04 2012-07-20 Snecma Soufflante de turbomachine
FR2968363B1 (fr) 2010-12-03 2014-12-05 Snecma Rotor de turbomachine avec une cale anti-usure entre un disque et un anneau
EP2546465A1 (en) 2011-07-14 2013-01-16 Siemens Aktiengesellschaft Blade root, corresponding blade, rotor disc, and turbomachine assembly
JP2013249756A (ja) * 2012-05-31 2013-12-12 Hitachi Ltd 圧縮機
EP2971568B1 (en) * 2013-03-15 2021-11-03 Raytheon Technologies Corporation Flap seal for a fan of a gas turbine engine
CA2918320C (en) 2013-07-26 2018-05-15 Mra Systems, Inc. Aircraft engine pylon
FR3014151B1 (fr) * 2013-11-29 2015-12-04 Snecma Soufflante, en particulier pour une turbomachine
FR3064667B1 (fr) * 2017-03-31 2020-05-15 Safran Aircraft Engines Dispositif de refroidissement d'un rotor de turbomachine
CN107100894A (zh) * 2017-07-05 2017-08-29 陕西金翼通风科技有限公司 一种通风机用叶片、叶轮及叶轮的安装方法
US10830048B2 (en) 2019-02-01 2020-11-10 Raytheon Technologies Corporation Gas turbine rotor disk having scallop shield feature
EP3862571A1 (en) * 2020-02-06 2021-08-11 ABB Schweiz AG Fan, synchronous machine and method for producing a fan

Family Cites Families (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2965355A (en) * 1956-01-17 1960-12-20 United Aircraft Corp Turbine disc burst inhibitor
US4344740A (en) * 1979-09-28 1982-08-17 United Technologies Corporation Rotor assembly
US4453890A (en) * 1981-06-18 1984-06-12 General Electric Company Blading system for a gas turbine engine
FR2519072B1 (fr) * 1981-12-29 1986-05-30 Snecma Dispositif de retenue axiale et radiale d'aube de rotor de turboreacteur
FR2695433B1 (fr) * 1992-09-09 1994-10-21 Snecma Joint annulaire d'étanchéité disposé à une extrémité axiale d'un rotor et recouvrant des brochages d'aubes.
US5281098A (en) * 1992-10-28 1994-01-25 General Electric Company Single ring blade retaining assembly
US5443365A (en) * 1993-12-02 1995-08-22 General Electric Company Fan blade for blade-out protection
RU2173390C2 (ru) * 1996-06-21 2001-09-10 Сименс Акциенгезелльшафт Ротор для турбомашины с устанавливаемыми в пазы лопатками, а также лопатка для ротора
JP2000512707A (ja) * 1996-06-21 2000-09-26 シーメンス アクチエンゲゼルシヤフト 溝内に装着可能な翼を有するタービン機械のロータ及びロータの翼
US6183202B1 (en) * 1999-04-30 2001-02-06 General Electric Company Stress relieved blade support
GB9925261D0 (en) * 1999-10-27 1999-12-29 Rolls Royce Plc Locking devices
FR2803623B1 (fr) * 2000-01-06 2002-03-01 Snecma Moteurs Agencement de retenue axiale d'aubes dans un disque
US6481971B1 (en) * 2000-11-27 2002-11-19 General Electric Company Blade spacer
US6634863B1 (en) * 2000-11-27 2003-10-21 General Electric Company Circular arc multi-bore fan disk assembly
GB2380770B (en) * 2001-10-13 2005-09-07 Rolls Royce Plc Indentor arrangement
GB2409240B (en) * 2003-12-18 2007-04-11 Rolls Royce Plc A gas turbine rotor
JP2005273646A (ja) * 2004-02-25 2005-10-06 Mitsubishi Heavy Ind Ltd 動翼体及びこの動翼体を有する回転機械
EP1703079A1 (de) * 2005-08-26 2006-09-20 Siemens Aktiengesellschaft Rotationskörper zum Befestigen von Laufschaufeln einer Strömungsmaschine

Also Published As

Publication number Publication date
FR2911632B1 (fr) 2009-08-21
RU2008101906A (ru) 2009-07-27
JP5283388B2 (ja) 2013-09-04
CA2619299C (fr) 2015-06-09
US8246309B2 (en) 2012-08-21
FR2911632A1 (fr) 2008-07-25
RU2454572C2 (ru) 2012-06-27
US20080298972A1 (en) 2008-12-04
EP1950381A1 (fr) 2008-07-30
CA2619299A1 (fr) 2008-07-18
JP2008180219A (ja) 2008-08-07

Similar Documents

Publication Publication Date Title
EP1950381B1 (fr) Disque de rotor de soufflante de turbomachine
EP2060750B1 (fr) Etage de turbine ou de compresseur, en particulier de turbomachine
EP1970537B1 (fr) Soufflante de turbomachine
EP2366061B1 (fr) Roue de turbine avec système de rétention axiale des aubes
EP1970538B1 (fr) Disque de rotor d'une turbomachine
CA2802821C (fr) Secteur angulaire de redresseur pour compresseur de turbomachine
FR2875262A1 (fr) Procedes et dispositifs pour assembler des ensembles de rotors de moteurs a turbines a gaz
CA2726018C (fr) Rotor de soufflante pour une turbomachine
FR2981979A1 (fr) Roue de turbine pour une turbomachine
EP2071129A1 (fr) Distributeur sectorisé pour une turbomachine
EP1213483A1 (fr) Etage redresseur d'un compresseur
FR3075869A1 (fr) Roue mobile de turbine pour turbomachine d'aeronef, comprenant un anneau d'etancheite retenu radialement par des excroissances sur l'echasse des aubes
FR3021693B1 (fr) Plateforme pour roue aubagee
EP1630350B1 (fr) Aube de rotor d'un compresseur ou d'une turbine à gaz
EP4402347B1 (fr) Aube mobile pour turbine de turbomachine, comprenant une échasse équipée d'excroissances de retenue radiale de l'aube
CA2577502A1 (fr) Roue de rotor de turbomachine
EP3444439A1 (fr) Turbine pour turbomachine comportant des aubes comprenant un pied presentant une forme s'évansant axialement
FR3127018A1 (fr) Aube mobile pour turbine de turbomachine, présentant une conception améliorant l’étanchéité des cavités inter-aubes
FR3124215A1 (fr) Procede pour reengager des clinquants pour aubes de turbomachine et dispositif de reengagement d’un clinquant
FR3100836A1 (fr) Aubes mobiles pour turbine
FR3144849A1 (fr) Ensemble rotorique de soufflante pour turbomachine
FR3126447A1 (fr) Roue mobile de turbomachine comprenant une pièce de butée axiale pour amortisseur
FR2973453A1 (fr) Aube de soufflante de turbomachine
FR3037367A1 (fr) Ensemble rotatif de turbomachine a garnitures de protection bombees
FR2989992A1 (fr) Dispositif d'immobilisation en rotation d'un anneau de retention d'aubes sur un disque de rotor

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 IS IT LI LT LU LV MC MT NL PL PT RO SE SI SK TR

AX Request for extension of the european patent

Extension state: AL BA HR MK RS

17P Request for examination filed

Effective date: 20080918

AKX Designation fees paid

Designated state(s): DE FR GB

REG Reference to a national code

Ref country code: DE

Ref legal event code: R079

Ref document number: 602007045027

Country of ref document: DE

Free format text: PREVIOUS MAIN CLASS: F01D0005300000

Ipc: F04D0029320000

RIC1 Information provided on ipc code assigned before grant

Ipc: F04D 29/32 20060101AFI20150630BHEP

Ipc: F01D 5/30 20060101ALI20150630BHEP

Ipc: F01D 21/04 20060101ALI20150630BHEP

GRAP Despatch of communication of intention to grant a patent

Free format text: ORIGINAL CODE: EPIDOSNIGR1

INTG Intention to grant announced

Effective date: 20150828

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: 602007045027

Country of ref document: DE

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 10

REG Reference to a national code

Ref country code: DE

Ref legal event code: R097

Ref document number: 602007045027

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: 20161205

REG Reference to a national code

Ref country code: FR

Ref legal event code: CD

Owner name: SNECMA, FR

Effective date: 20170713

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 11

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

Ref country code: DE

Payment date: 20241121

Year of fee payment: 18

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

Ref country code: GB

Payment date: 20241122

Year of fee payment: 18

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

Ref country code: FR

Payment date: 20241121

Year of fee payment: 18