[go: up one dir, main page]
More Web Proxy on the site http://driver.im/

EP1950381A1 - Rotor disc for turbomachine fan - Google Patents

Rotor disc for turbomachine fan Download PDF

Info

Publication number
EP1950381A1
EP1950381A1 EP07291632A EP07291632A EP1950381A1 EP 1950381 A1 EP1950381 A1 EP 1950381A1 EP 07291632 A EP07291632 A EP 07291632A EP 07291632 A EP07291632 A EP 07291632A EP 1950381 A1 EP1950381 A1 EP 1950381A1
Authority
EP
European Patent Office
Prior art keywords
blade
cavities
disk
grooves
platforms
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.)
Granted
Application number
EP07291632A
Other languages
German (de)
French (fr)
Other versions
EP1950381B1 (en
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/en
Application granted granted Critical
Publication of EP1950381B1 publication Critical patent/EP1950381B1/en
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.
  • 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 invention aims in particular to provide a simple, economical and effective solution to these various problems.
  • a fan rotor disk in a turbomachine comprising at the periphery substantially axial grooves for mounting and retaining blade roots having hooks at their downstream ends, deformable zones formed by cavities being located at the downstream end of the grooves, characterized in that the cavities are formed in clamping flanges of inter-blade platforms.
  • 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 out on the notch 26.
  • the axial machining 38 makes it possible 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)

Abstract

The disc has axial grooves (22) for mounting and retaining the vane roots, and deformable zones (34) radially located outside the grooves, where the zones are placed at a rear end of the grooves. The zones having cavities are formed in hooking flanges (36) of inter-vane platforms, and are placed at the level of the rear notches of the vane roots, where the platforms orient air flow at inlet of a turbomachine.

Description

La présente invention concerne un disque de rotor de soufflante pour turbomachine, telle en particulier qu'un turboréacteur d'avion.The present invention relates to a fan rotor disk for a turbomachine, such in particular as an airplane turbojet engine.

De façon connue, un disque de rotor de soufflante comprend une pluralité d'aubes montées à sa périphérie et séparées entres elles par des plates-formes fixées sur des brides du disque. Chaque aube est formée d'une pale raccordée à un pied d'aube par l'intermédiaire d'une échasse. Les pieds d'aubes sont engagés dans des rainures sensiblement axiales, formées en périphérie du disque et y sont maintenus radialement par coopération de formes, les pieds d'aube étant par exemple à section transversale en queue d'aronde ou analogue.In known manner, 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.

Lors du fonctionnement de la turbomachine, la rupture de la liaison d'une aube avec le disque peut entraîner la destruction des aubes voisines et des plates-formes attenantes. En effet, en cas de perte d'aube de soufflante, celle-ci vient en appui sur l'aube voisine, et l'effort résultant appliqué à cette aube se traduit notamment par une contrainte axiale dirigée de l'aval vers l'amont du fait du calage angulaire de la pale par rapport à la rainure, ce qui a tendance à faire basculer l'aube vers l'amont et à générer une forte contrainte au niveau de la liaison arrière entre le pied d'aube et le disque. Une cassure du pied d'aube ou d'une dent du disque peut ainsi survenir, conduisant à une réaction en chaîne pouvant détruire toutes les aubes de la soufflante ainsi que les plates-formes et endommageant fortement la turbomachine.During the operation of the turbomachine, 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.

Dans certains types d'aubes, le pied d'aube qui est engagé dans la rainure, est raccordé en aval à un crochet. Des encoches formées radialement de part et d'autre de chaque crochet coopèrent avec un flasque annulaire afin d'assurer le maintien axial des aubes lorsqu'elles sont positionnées dans les rainures du disque. En cas de perte d'aube, ce mode de fixation génère une forte contrainte au niveau de la zone de raccordement de l'échasse avec le crochet et au niveau du raccordement de l'encoche avec le crochet. Comme précédemment, cette contrainte peut engendrer une cassure, au niveau du crochet de l'aube ou au niveau du disque, et entraîner une destruction en chaîne des aubes et des plates-formes .In certain types of blades, 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. In case of dawn loss, 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.

Dans la technique actuelle, une rainure axiale d'environ 10 mm de long, débouchant sur l'encoche, est usinée de chaque côté du pied d'aube, pour limiter la contrainte appliquée au niveau de la zone de raccordement de l'échasse avec le crochet et au niveau de la zone de raccordement entre l'encoche et le crochet, en orientant les efforts en amont de l'usinage. Cette rainure, si elle permet de limiter les efforts au niveau du crochet, présente néanmoins l'inconvénient de générer un pic de contrainte à son extrémité amont ce qui engendre une usure importante du pied d'aube et du disque et limite ainsi leur durée de vie. Plusieurs solutions ont été envisagées pour limiter l'usure de ces pièces et ont consisté à former un détalonnage à l'extrémité amont de l'usinage, ou à placer un clinquant entre l'aube et le disque. Cependant, ces moyens ne permettent pas de résoudre de manière satisfaisante le problème d'usure tout en limitant la contrainte appliquée au crochet de l'aube et transmise aux plates-formes.In the present technique, 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.

L'invention a notamment pour but d'apporter une solution simple, économique et efficace à ces différents problèmes.The invention aims in particular to provide a simple, economical and effective solution to these various problems.

Elle propose à cet effet un disque de rotor de soufflante dans une turbomachine, comprenant en périphérie des rainures sensiblement axiales de montage et de retenue de pieds d'aubes comportant des crochets à leurs extrémités aval, des zones déformables formées par des cavités étant situées à l'extrémité aval des rainures, caractérisé en ce que les cavités sont formées dans des brides d'accrochage de plates-formes inter-aubes.It proposes for this purpose a fan rotor disk in a turbomachine, comprising at the periphery substantially axial grooves for mounting and retaining blade roots having hooks at their downstream ends, deformable zones formed by cavities being located at the downstream end of the grooves, characterized in that the cavities are formed in clamping flanges of inter-blade platforms.

En cas de perte d'aube, les contraintes exercées par les pieds d'aube sur le disque sont maximales à l'extrémité aval du disque et induisent une déformation locale plastique des cavités situées au niveau des brides d'accrochage des plates-formes inter-aubes, ce qui limite le niveau de contrainte appliqué au disque ainsi qu'aux plates-formes inter-aubes. Les aubes et les plates-formes peuvent ainsi être maintenues en position durant le temps nécessaire à l'arrêt du moteur, ce qui évite ainsi d'importants endommagements de la turbomachine.In the event of dawn loss, the stresses exerted by the blade roots on the disk are maximum at the downstream end of the disk and induce a plastic local deformation of the cavities located at the inter-platform hooking flanges. -aubes, which limits the level of constraint applied to the disk as well as to the platforms inter-blade. The blades and the platforms can thus be held in position during the time necessary to stop the engine, thereby avoiding significant damage to the turbomachine.

Les aubes du disque de rotor selon l'invention ne nécessitent plus un usinage axial permettant la déviation des efforts. On supprime ainsi les phénomènes d'usure du disque et de l'aube dus à cet usinage tout en limitant les contraintes appliquées aux crochets et transmise aux plates-formes, grâce aux cavités réalisées dans les brides d'accrochage de plates-formes inter-aubes.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.

Selon une autre caractéristique de l'invention, les cavités sont formées par usinage.According to another characteristic of the invention, the cavities are formed by machining.

Avantageusement, les cavités sont orientées axialement et sont de forme tubulaire à fond fermé.Advantageously, the cavities are oriented axially and are of closed bottom tubular shape.

Dans une réalisation de l'invention, les cavités sont formées par perçage ou fraisage.In one embodiment of the invention, the cavities are formed by drilling or milling.

Dans une autre variante de l'invention, les cavités sont ouvertes latéralement et débouchent à l'intérieur des rainures.In another variant of the invention, the cavities are open laterally and open into the grooves.

L'invention concerne également une turbomachine, telle qu'un turboréacteur d'avion, caractérisée en ce qu'elle comprend un disque de rotor de soufflante du type décrit ci-dessus.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.

D'autres avantages et caractéristiques de l'invention apparaîtront à la lecture de la description suivante faite à titre d'exemple non limitatif et en référence aux dessins annexés dans lesquels :

  • la figure 1 est une vue partielle en perspective d'un disque selon l'invention ;
  • la figure 2 est une vue en perspective de la partie aval d'un pied d'aube de soufflante selon la technique antérieure ;
  • la figure 3 est une vue schématique en perspective d'un premier mode de réalisation d'un disque de rotor selon l'invention ;
  • la figure 4 est une vue schématique en perspective d'un deuxième mode de réalisation d'un disque de rotor selon l'invention ;
  • la figure 5 est une vue schématique en perspective d'un troisième mode de réalisation d'un disque de rotor selon l'invention.
Other advantages and characteristics of the invention will appear on reading the following description given by way of nonlimiting example and with reference to the appended drawings in which:
  • the figure 1 is a partial perspective view of a disk according to the invention;
  • the figure 2 is a perspective view of the downstream portion of a blower blade root according to the prior art;
  • the figure 3 is a schematic perspective view of a first embodiment of a rotor disk according to the invention;
  • the figure 4 is a schematic perspective view of a second embodiment of a rotor disk according to the invention;
  • the figure 5 is a schematic perspective view of a third embodiment of a rotor disk according to the invention.

On se réfère tout d'abord à la figure 1, représentant un disque 10 de soufflante portant une aube 12 ainsi qu'à la figure 2 qui représente la partie aval radialement interne d'une aube selon la technique antérieure.We first refer to the figure 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.

Une aube est formée d'une pale 14 raccordée à un pied d'aube 20 par l'intermédiaire d'une échasse 18. Le disque 10 comporte une pluralité de rainures 22 sensiblement axiales réparties régulièrement à sa périphérie externe et dans lesquelles sont engagées les aubes 12. Des plates-formes (non représentées) sont disposées entre les aubes et servent à orienter le flux d'air en entrée de la turbomachine. Le pied d'aube 20 en forme de queue d'aronde ou analogue coopère avec la rainure 22 afin d'assurer la retenue radiale de l'aube (12) sur le disque 10 de rotor. Dans le prolongement aval du pied d'aube 20 du disque 10 est formé un crochet 24 comprenant une encoche 26 radiale sur chacune de ses faces latérales. Ces encoches coopèrent avec un flasque annulaire 28 pour bloquer axialement le pied 20 de l'aube 12 dans la rainure 22 du disque 10.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. In the downstream extension of the blade root 20 of the disk 10 is formed 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.

Lors du fonctionnement de la turbomachine, les zones de raccordement échasse/crochet 30 et encoche/crochet 32 sont fortement sollicitées. En cas de perte d'aube, le contact radial de l'aube désolidarisée du disque avec l'aube voisine se traduit du fait de la fixation de l'aube dans une rainure par une contrainte supplémentaire dans les zones de raccordement échasse/crochet 30 et encoche/crochet 32. Des lors, la contrainte appliquée à l'arrière de l'aube fragilise le crochet 24, ce qui peut entraîner sa rupture. Une telle contrainte peut également endommager le disque et donc les plates-formes inter-aubes qui y sont fixées. La rupture de la liaison avec le disque d'une deuxième aube peut entraîner une réaction en chaîne conduisant à la destruction complète des aubes de soufflante et des plates-formes attenantes, conduisant à un important endommagement de la turbomachine. Il est donc impératif de maintenir les aubes en position dans leurs rainures ainsi que les plates-formes sur les brides d'accrochage du disque en cas de perte d'aubes.During operation of the turbomachine, the hook / hook 30 and notch / hook connection zones 32 are strongly stressed. In the event of dawn loss, 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.

Dans la technique connue, représentée en figure 2, un usinage axial 38 est réalisé de chaque côté du crochet 24, et débouche sur l'encoche 26. L'usinage axial 38 permet de déporter les efforts, représentés en flèches pointillées, au-delà de l'usinage ce qui réduit les contraintes appliquées au crochet, les efforts en l'absence d'usinage étant représentés en flèches pleines. Les contraintes appliquées au crochet sont ainsi limitées et l'aube bénéficie d'une meilleure tenue. Cependant, ce type de solution n'est pas satisfaisant puisqu'une forte contrainte est générée à l'extrémité amont de l'usinage 38, ce qui engendre une usure importante du pied d'aube et du disque.In the known art, represented in figure 2 , an axial machining 38 is performed on each side of the hook 24, and opens out on the notch 26. The axial machining 38 makes it possible 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. However, 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.

L'invention propose, pour supprimer ce phénomène d'usure tout en limitant la contrainte appliquée à la liaison de l'aube avec le disque et transmise aux plates-formes, de former des zones déformables 34 dans le disque 10 situées radialement à l'extérieur des rainures 22, au niveau des crochets des pieds d'aube.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.

Comme représenté aux figures 3, 4 et 5, des zones déformables 34 sont formées par des cavités 34 réalisées dans des brides d'accrochage 36 de plates-formes inter-aubes (non représentées), et sont fixées sur des brides 36 s'étendant sensiblement dans le prolongement des parois latérales des rainures 22 (figure 3 à 5)As represented in Figures 3, 4 and 5 , 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 )

En figures 3 et 4 sont représentés deux premiers modes de réalisation de l'invention dans lesquels les cavités 34 sont orientées axialement et sont de forme tubulaires à fond fermé.In Figures 3 and 4 two first embodiments of the invention are shown in which the cavities 34 are oriented axially and are of closed-bottom tubular shape.

Dans un troisième mode de réalisation de l'invention, représenté en figure 5, les cavités 34 sont ouvertes latéralement et débouchent à l'intérieur des rainures.In a third embodiment of the invention, shown in figure 5 the cavities 34 are open laterally and open into the grooves.

Dans ces différents modes de réalisation, la cavité a par exemple un diamètre de l'ordre de 6 à 9 mm, l'épaisseur de la paroi de la cavité est comprise entre 0 et 3 mm, la profondeur étant d'environ 20 mm. Ces valeurs sont données à titre indicatif pour un disque 10 de rotor de diamètre extérieur de l'ordre de 200 mm.In these different embodiments, 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. These values are given as an indication for a rotor disc 10 with an outside diameter of the order of 200 mm.

Ces cavités peuvent être réalisées par des techniques d'usinage simples et rapides telles que le perçage ou bien le fraisage.These cavities can be made by simple and fast machining techniques such as drilling or milling.

L'intégration de cavités 34 dans les brides 36 d'accrochage des plates-formes inter-aubes autorise la déformation plastique de ces cavités en cas de perte d'aube. Les efforts en sortie de portée d'aube sont orientés vers les cavités 34. Ainsi la contrainte appliquée au crochet arrière est plus faible, ce qui évite la rupture du crochet et permet à l'aube de rester en position dans sa rainure et aux plates-formes attenantes de rester fixées sur les brides 36 du disque 10 jusqu'à l'arrêt de la turbomachine. De plus, en fonctionnement normal, la durée de vie n'est plus limitée par les phénomènes d'usures dus à un usinage axial dans le pied d'aube 20, celui-ci n'étant plus nécessaire.The integration of 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. Thus 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. In addition, in normal operation, 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.

Si l'invention précédemment décrite est particulièrement intéressante dans le cas d'une utilisation combinée avec des aubes 12 à crochets 24, elle n'est cependant pas limitée à ce type d'application et peut être utilisée avec tous les autres types d'aubes 12 de soufflante.If the invention described above is particularly advantageous in the case of a combined use with blades 12 hooks 24, it is however not limited to this type of application and can be used with all other types of blades 12 of blower.

Claims (6)

Disque (10) de rotor de soufflante dans une turbomachine, comprenant en périphérie des rainures (22) sensiblement axiales de montage et de retenue de pieds d'aubes (20) comportant des crochets à leurs extrémités aval, des zones déformables formées par des cavités (34) étant situées à l'extrémité aval des rainures (22), caractérisé en ce que les cavités (34) sont formées dans des brides d'accrochage (36) de plates-formes inter-aubes.Disk (10) for a fan rotor in a turbomachine, comprising on the periphery substantially axial grooves (22) for mounting and retaining blade roots (20) comprising hooks at their downstream ends, deformable zones formed by cavities (34) being located at the downstream end of the grooves (22), characterized in that the cavities (34) are formed in latching clamps (36) of inter-blade platforms. Disque selon la revendication 1, caractérisé en ce que les cavités (34) sont formées par usinage.Disk according to claim 1, characterized in that the cavities (34) are formed by machining. Disque selon la revendication 1 ou 2, caractérisé en ce que les cavités (34) sont orientées axialement et sont de forme tubulaire à fond fermé.Disk according to claim 1 or 2, characterized in that the cavities (34) are oriented axially and are of closed bottom tubular form. Disque selon l'une des revendications 1 à 3, caractérisé en ce que les cavités (34) sont formées par perçage ou par fraisage.Disk according to one of claims 1 to 3, characterized in that the cavities (34) are formed by drilling or milling. Disque selon l'une des revendications 1 à 4, caractérisé en ce que les cavités (34) sont ouvertes latéralement et débouchent à l'intérieur des rainures (22).Disk according to one of claims 1 to 4, characterized in that the cavities (34) are open laterally and open into the grooves (22). Turbomachine, telle qu'un turboréacteur d'avion, caractérisée en ce qu'elle comprend un disque (10) de rotor de soufflante selon l'une des revendications précédentes.Turbomachine, such as an aircraft turbojet engine, characterized in that it comprises a fan rotor disc (10) according to one of the preceding claims.
EP07291632.3A 2007-01-18 2007-12-27 Rotor disc for turbomachine fan Active EP1950381B1 (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0700326A FR2911632B1 (en) 2007-01-18 2007-01-18 ROTOR DISC OF TURBOMACHINE BLOWER

Publications (2)

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

Family

ID=38421439

Family Applications (1)

Application Number Title Priority Date Filing Date
EP07291632.3A Active EP1950381B1 (en) 2007-01-18 2007-12-27 Rotor disc for turbomachine fan

Country Status (6)

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

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2010088882A3 (en) * 2009-02-04 2011-02-24 Mtu Aero Engines Gmbh Integrally bladed rotor disk for a turbine
FR2968363A1 (en) * 2010-12-03 2012-06-08 Snecma Rotor for use in turbojet engine e.g. turbofan engine, has annular wedge arranged and interposed between disk and ring and forming axial support surface against which paddles placed in slots are supported
CN102753788A (en) * 2010-02-04 2012-10-24 斯奈克玛 Turbine engine air blower

Families Citing this family (12)

* 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
US8485784B2 (en) * 2009-07-14 2013-07-16 General Electric Company Turbine bucket lockwire rotation prevention
EP2299056A1 (en) * 2009-09-02 2011-03-23 Siemens Aktiengesellschaft Cooling of a gas turbine component shaped as a rotor disc or as a blade
EP2546465A1 (en) 2011-07-14 2013-01-16 Siemens Aktiengesellschaft Blade root, corresponding blade, rotor disc, and turbomachine assembly
JP2013249756A (en) * 2012-05-31 2013-12-12 Hitachi Ltd Compressor
EP2971568B1 (en) * 2013-03-15 2021-11-03 Raytheon Technologies Corporation Flap seal for a fan of a gas turbine engine
CN105392700B (en) 2013-07-26 2018-12-18 Mra系统有限责任公司 Aircraft engine hanger
FR3014151B1 (en) * 2013-11-29 2015-12-04 Snecma BLOWER, ESPECIALLY FOR A TURBOMACHINE
FR3064667B1 (en) * 2017-03-31 2020-05-15 Safran Aircraft Engines DEVICE FOR COOLING A TURBOMACHINE ROTOR
CN107100894A (en) * 2017-07-05 2017-08-29 陕西金翼通风科技有限公司 A kind of installation method of ventilation blower blade, impeller and impeller
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

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2100808A (en) 1981-06-18 1983-01-06 Gen Electric Blade root modification for a gas turbine engine
EP1048821A2 (en) * 1999-04-30 2000-11-02 General Electric Company Stress relieved blade root
GB2380770A (en) 2001-10-13 2003-04-16 Rolls Royce Plc Stress-reducing indentor profile for gas turbine engine blade mountings and other applications
US6634863B1 (en) 2000-11-27 2003-10-21 General Electric Company Circular arc multi-bore fan disk assembly
EP1703079A1 (en) * 2005-08-26 2006-09-20 Siemens Aktiengesellschaft Rotational solid for fixing of blades of a turbo-machine

Family Cites Families (14)

* 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
SU387128A1 (en) * 1971-05-24 1973-06-21 WORKING WHEEL TURBO MOBILE
US4344740A (en) * 1979-09-28 1982-08-17 United Technologies Corporation Rotor assembly
FR2519072B1 (en) * 1981-12-29 1986-05-30 Snecma DEVICE FOR AXIAL AND RADIAL RETENTION OF A TURBO JET ROTOR BLADE
FR2695433B1 (en) * 1992-09-09 1994-10-21 Snecma Annular seal placed at an axial end of a rotor and covering blade pinouts.
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 (en) * 1996-06-21 2001-09-10 Сименс Акциенгезелльшафт Turbo-machine rotor accommodating blades in its slots and rotor blades
WO1997049921A1 (en) * 1996-06-21 1997-12-31 Siemens Aktiengesellschaft Rotor for a turbomachine with blades insertable into grooves and blades for a rotor
GB9925261D0 (en) * 1999-10-27 1999-12-29 Rolls Royce Plc Locking devices
FR2803623B1 (en) * 2000-01-06 2002-03-01 Snecma Moteurs ARRANGEMENT FOR AXIAL RETENTION OF BLADES IN A DISC
US6481971B1 (en) * 2000-11-27 2002-11-19 General Electric Company Blade spacer
GB2409240B (en) * 2003-12-18 2007-04-11 Rolls Royce Plc A gas turbine rotor
JP2005273646A (en) * 2004-02-25 2005-10-06 Mitsubishi Heavy Ind Ltd Moving blade element and rotary machine having the moving blade element

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2100808A (en) 1981-06-18 1983-01-06 Gen Electric Blade root modification for a gas turbine engine
EP1048821A2 (en) * 1999-04-30 2000-11-02 General Electric Company Stress relieved blade root
US6634863B1 (en) 2000-11-27 2003-10-21 General Electric Company Circular arc multi-bore fan disk assembly
GB2380770A (en) 2001-10-13 2003-04-16 Rolls Royce Plc Stress-reducing indentor profile for gas turbine engine blade mountings and other applications
EP1703079A1 (en) * 2005-08-26 2006-09-20 Siemens Aktiengesellschaft Rotational solid for fixing of blades of a turbo-machine

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2010088882A3 (en) * 2009-02-04 2011-02-24 Mtu Aero Engines Gmbh Integrally bladed rotor disk for a turbine
US8821122B2 (en) 2009-02-04 2014-09-02 Mtu Aero Engines Gmbh Integrally bladed rotor disk for a turbine
CN102753788A (en) * 2010-02-04 2012-10-24 斯奈克玛 Turbine engine air blower
CN102753788B (en) * 2010-02-04 2015-02-11 斯奈克玛 Turbine engine air blower
FR2968363A1 (en) * 2010-12-03 2012-06-08 Snecma Rotor for use in turbojet engine e.g. turbofan engine, has annular wedge arranged and interposed between disk and ring and forming axial support surface against which paddles placed in slots are supported
US8911212B2 (en) 2010-12-03 2014-12-16 Snecma Turbomachine rotor with anti-wear shim between a disk and an annulus

Also Published As

Publication number Publication date
RU2008101906A (en) 2009-07-27
JP2008180219A (en) 2008-08-07
US8246309B2 (en) 2012-08-21
CA2619299A1 (en) 2008-07-18
US20080298972A1 (en) 2008-12-04
CA2619299C (en) 2015-06-09
JP5283388B2 (en) 2013-09-04
RU2454572C2 (en) 2012-06-27
EP1950381B1 (en) 2016-03-02
FR2911632A1 (en) 2008-07-25
FR2911632B1 (en) 2009-08-21

Similar Documents

Publication Publication Date Title
EP1950381B1 (en) Rotor disc for turbomachine fan
EP2060750B1 (en) Stage of a turbine or compressor, in particular of a turbomachine
CA2625317C (en) Turbomachine rotor disk
EP2366061B1 (en) Turbine wheel with an axial retention system for vanes
CA2521265C (en) Excess turbine speed limitation device in a turbine engine
CA2625319A1 (en) Turbomachine ventilator
CA2598532C (en) Turbomachine rotor blade
FR2875262A1 (en) METHODS AND DEVICES FOR ASSEMBLING ROTOR ASSEMBLIES OF GAS TURBINE ENGINES
FR2981979A1 (en) TURBINE WHEEL FOR A TURBOMACHINE
EP2071129A1 (en) Sectorised distributor for a turbomachine
EP1760259A2 (en) Locking apparatus for an axial retention ring of a blade
EP2300685A1 (en) Turbomachine fan rotor
EP1213483A1 (en) Compressor stator stage
FR3075869A1 (en) MOBILE TURBINE WHEEL FOR AIRCRAFT TURBOMACHINE, COMPRISING A SEAL RING RADIALLY RETAINED BY INCREASES ON THE ECHASSE DES AUBES
FR3099520A1 (en) Turbomachine wheel
EP1630350B1 (en) Rotor blade of a compressor or a gas turbine
CA2577502A1 (en) Turbine engine rotor wheel
WO2023041868A1 (en) Moving blade for a turbine of a turbine engine, comprising a stilt equipped with projections for radially retaining the blade
EP3444439A1 (en) Turbine for turbine engine comprising blades with a root having an exapnding form in axial direction
EP1630351A1 (en) Blade for a compressor or a gas turbine
FR3100836A1 (en) MOBILE BLADES FOR TURBINE
FR3127018A1 (en) Moving blade for a turbomachine turbine, having a design improving the sealing of the inter-blade cavities
FR3144849A1 (en) FAN ROTOR ASSEMBLY FOR TURBOMACHINE
FR3026794A1 (en) ROTARY ASSEMBLY FOR TURBOMACHINE AND FASTENING PION FOR THIS SET
FR3144848A1 (en) FAN ROTOR ASSEMBLY FOR 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 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: GB

Payment date: 20231121

Year of fee payment: 17

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

Ref country code: FR

Payment date: 20231122

Year of fee payment: 17

Ref country code: DE

Payment date: 20231121

Year of fee payment: 17