EP3759319B1 - Assembly for a turbomachine - Google Patents

Assembly for a turbomachine Download PDF

Info

Publication number
EP3759319B1
EP3759319B1 EP19717517.7A EP19717517A EP3759319B1 EP 3759319 B1 EP3759319 B1 EP 3759319B1 EP 19717517 A EP19717517 A EP 19717517A EP 3759319 B1 EP3759319 B1 EP 3759319B1
Authority
EP
European Patent Office
Prior art keywords
radially
support
turbomachine
turbine
annular channel
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
EP19717517.7A
Other languages
German (de)
French (fr)
Other versions
EP3759319A1 (en
Inventor
Baptiste HALLOUIN
Alexandre Montpellaz
Sylvain Pierre Votie
Fabrice Iparaguirre
Yann DANIS
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 Helicopter Engines SAS
Original Assignee
Safran Helicopter Engines 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 Safran Helicopter Engines SAS filed Critical Safran Helicopter Engines SAS
Priority to PL19717517T priority Critical patent/PL3759319T3/en
Publication of EP3759319A1 publication Critical patent/EP3759319A1/en
Application granted granted Critical
Publication of EP3759319B1 publication Critical patent/EP3759319B1/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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/16—Arrangement of bearings; Supporting or mounting bearings in casings
    • F01D25/162—Bearing supports
    • F01D25/164—Flexible supports; Vibration damping means associated with the bearing
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • F01D9/00—Stators
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/18—Lubricating arrangements
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F05D2220/323—Application in turbines in gas turbines for aircraft propulsion, e.g. jet engines
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/60—Assembly methods
    • 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
    • F05D2240/20—Rotors
    • F05D2240/24—Rotors for turbines

Definitions

  • the present invention relates to an assembly for a turbomachine, such as, for example, an aircraft turbojet or turboprop engine.
  • the figure 1 illustrates part of a turbomachine 1 according to a first embodiment in accordance with the prior art.
  • upstream and downstream are defined with respect to the direction of circulation of the gases within the turbomachine 1.
  • the turbomachine 1 comprises an upstream turbine 2 and a downstream turbine 3.
  • the upstream turbine 2 is for example a high-pressure turbine and the downstream turbine 3 is for example a low-pressure turbine or a free turbine.
  • Each turbine 2, 3 comprises a rotor comprising blades 4.
  • the turbine engine 1 further comprises a radially internal shaft 5, extending along the axis A of the turbine engine 1.
  • the turbomachine 1 further comprises an annular channel 6 intended to form a flow path for the gas flow between two turbine stages 2, 3 of the turbomachine 1, said channel 6 being delimited by a radially internal annular wall 7 and a wall radially outer ring finger 8.
  • a radially outer support 9 connects the outer annular wall and a casing 10 of the turbine.
  • the external support 9 comprises a flexible or elastically deformable zone 11, able to allow the radial and/or axial displacement of the external annular wall 8 with respect to the casing 10.
  • a radially internal support 12 extends radially inwards from the radially internal wall 7.
  • the radially internal part 13 of the internal support 11 surrounds two bearings 14 mounted around the shaft 5.
  • the internal support 12 further comprises a zone 15 flexible or elastic deformable, capable of allowing the radial and/or axial displacement of the internal annular wall 7 with respect to the bearings 14 and to the shaft 5.
  • the assembly formed by the annular channel 6 and the internal and external supports 9, 12 is made in one piece, for example by foundry.
  • the internal and external annular walls 7, 8 of the annular channel 6 are subjected to high temperatures, while the internal 12 and external 9 supports can be subjected to lower temperatures.
  • the temperature difference is in particular very large during the so-called transition phase, when the turbine engine is started. This temperature difference generates differential expansions between the different parts of the same assembly.
  • the flexible zones 11, 15 of the supports 9, 12 make it possible to compensate for such differential expansions, by allowing radial and/or axial displacement of the internal and external annular walls 7, 8 of the annular channel 6 with respect to the other parts of the assembly. .
  • the supports 9, 12 perform a so-called structural function since their particular function is to support radially the shaft 5, that is to say to link it to the casing 10, and to avoid the radial displacement of the shaft 5, in particular under load.
  • the assembly is made, for example, of a nickel-based alloy of the Inconel 738 type, such a material being expensive and not being repairable by reloading the material by welding.
  • FIG. 2 A second embodiment known from the prior art is shown in figure 2 .
  • the assembly comprises an annular channel 6 intended to form a flow path for a flow of gas between the two turbine stages 2, 3 of the turbomachine 1, said channel 6 being delimited by a wall radially inner annular wall 7 and a radially outer annular wall 8, said walls 7, 8 being connected by hollow arms 16 extending radially.
  • the assembly further comprises a support 17, separate from the annular channel 6, and comprising a radially outer annular part 9, located radially outside the outer annular wall 8 of the annular channel 6, and a radially inner annular part 12, located radially inside the internal annular wall 7 of the annular channel 6, the external 9 and internal 12 parts of the support 17 being connected by connecting parts 18 extending radially, each connecting part 18 passing through a hollow arm 16 of the annular canal 6.
  • the invention aims to remedy these drawbacks, in a simple, reliable and inexpensive manner.
  • the assembly can thus be produced in a single piece, for example by additive manufacturing or by foundry, which makes it possible to reduce the manufacturing costs.
  • the annular channel and the support form two separate parts, so as to avoid conduction or thermal bridges by contact between said parts.
  • the breakable part can be sized to break when the shear stresses in the connecting partition, at the level of the breakable part, are greater than 200 MPa.
  • the aforementioned stress value is for example the value when the connecting partition is at a temperature between 500 and 900° C., this value being able to change with the temperature.
  • the assembly is made in one piece from a nickel-based alloy, for example from an alloy of the C263 type.
  • the alloy used can be hardfaced by welding. This is in particular the case of an alloy of the C263 type.
  • the breakable part can be formed by a thinned zone of the connecting partition.
  • the breakable part may include removals of material, such as holes or localized recessed areas, for example.
  • connection parts of the support may comprise an internal duct allowing the supply of a lubricating fluid from a zone located radially outside the annular channel to a zone located inside the annular channel .
  • the radially inner part of the support may be intended to support at least one bearing.
  • the duct can thus allow the lubrication of said bearing.
  • the lubricating fluid is for example grease or oil.
  • the radially internal part and/or the radially external part of the support may comprise at least one flexible zone allowing radial deformation of said radially internal or external part.
  • the radially inner part and/or the radially outer part of the support can comprise a radially fixed peripheral part, connected to each connecting part by the corresponding flexible zone.
  • the flexible zone can be formed by elastically deformable tabs or pins.
  • Said tabs or pins can be oriented obliquely, that is to say can form a non-zero angle with the axial direction and with the radial direction.
  • Said angle with the axial direction is for example between 30 and 60°, for example of the order of 45°.
  • the invention relates to a turbomachine, such as for example a turbojet or a turboprop, comprising an upstream turbine, for example a high-pressure turbine, and a downstream turbine, for example a low-pressure turbine or a free turbine, said turbines each comprising a rotor, the turbomachine comprising a radially internal shaft, characterized in that it comprises an assembly of the aforementioned type, the annular channel forming a gas flow stream between the upstream turbine and the downstream turbine, the radially internal part of the support supporting at least one bearing serving to guide the shaft, the radially outer part of the support being fixed to a fixed part of the turbomachine, for example a turbine casing.
  • the temperature differential allowing a break in the breakable zone is for example between 200 and 500° C.
  • the breakable part can be broken cold, that is to say without heating part of the assembly, before mounting the annular channel and the support in the turbomachine.
  • the breakable part can be broken cold, that is to say without heating part of the assembly, after mounting the annular channel and the support in the turbomachine.
  • a stress can be generated mechanically at the level of the connecting partition, for example by an operator, in particular by applying a shock or a sufficient force to said partition.
  • the picture 3 represents a part of a turbomachine 1 according to one embodiment of the invention.
  • This comprises an upstream turbine 2 and a downstream turbine 3.
  • the upstream turbine 2 is for example a high-pressure turbine and the downstream turbine 3 is for example a low-pressure turbine or a free turbine.
  • Each turbine 2, 3 comprises a rotor comprising blades 4.
  • the turbomachine 1 also comprises a radially internal shaft 5, extending along the axis A of the turbomachine.
  • the turbomachine 1 further comprises an assembly comprising an annular channel 6 intended to form a flow path for a flow of gas between the two stages 2, 3 of the turbine of the turbomachine 1, said channel 6 being delimited by an annular wall radially inner 7 and a radially outer annular wall 8, said walls 7, 8 being connected by hollow arms 16 extending radially.
  • the whole also visible at the figure 4 , further comprises a support 17 comprising a radially outer annular part 9, located radially outside the outer annular wall 8 of the annular channel 6, and a radially inner annular part 12, located radially inside the annular wall 7 of the annular channel 6, the outer and inner parts 9, 12 of the support 17 being connected by connecting parts 18 extending radially, each connecting part 18 passing through one of the hollow arms 16 of the annular channel 6.
  • hollow arm 16 and the connecting parts 18 are regularly distributed over the periphery.
  • the radially internal part 12 and the radially external part 9 of the support 17 each comprise a flexible zone 11, 15 allowing radial deformation of said radially internal or external part 12, 9.
  • the radially inner part 12 comprises a radially outer annular flange 19, extending radially, and fixed to the casing 10 by means of screws or rivets for example.
  • Said flange 19 is connected to each connecting part 18 by the corresponding flexible zone 11.
  • This flexible zone 11 is formed by lugs or by elastically deformable pins 20.
  • Said tabs or pins 20 can be oriented obliquely, that is to say can form a non-zero angle with the axial direction and with the radial direction.
  • Said angle with the axial direction is for example between 30 and 60°, for example of the order of 45°.
  • the radially internal part 12 of the support 17 comprises annular parts 13a, 13b extending axially, each intended to surround one of the bearings 14.
  • Each annular part 13a, 13b is connected to the connecting parts 18 by zones flexible 15a, 15b oblique or tapered.
  • Each flexible oblique or tapered zone 15a, 15b forms a non-zero angle with the axial and radial directions.
  • At least one of the connecting parts 18 of the support 17 comprises an internal duct 21 allowing the supply of a lubricating fluid from a zone located radially outside the annular channel 6 into a area located opposite the bearings 14.
  • the lubricating fluid is for example grease or oil.
  • Each connecting part can have two rectilinear parts 18a, 18b forming an angle with respect to each other.
  • other shapes are also possible.
  • At least one of the connecting parts 18 and the corresponding hollow arm 16 are connected to each other by at least one connecting partition 22, said connecting partition 22 comprising a breakable part 23 capable of breaking when the mechanical stresses in said connecting partition 22 are greater than a determined value.
  • the connecting part 18 is not in contact with the surface of the connecting arm 16, so as to limit heat exchange.
  • the breakable part 23 can be sized to break when the shear stresses in the connecting partition 22, at the level of the breakable part 23, are greater than 200 MPa. This value can change with temperature and can for example be set at a temperature between 500°C and 900°C.
  • the assembly formed by the channel 6 and the support 17 can thus be produced in a single piece, for example by additive manufacturing or by foundry, which makes it possible to reduce the manufacturing costs.
  • the annular channel 6 and the support 17 form two separate parts, so as to avoid conduction or thermal bridges by contact between said parts 6, 17.
  • the assembly is made in one piece from a nickel-based alloy, for example from an alloy of the C263 type.
  • the breakable part 23 of the connecting partition 22 is formed by a thinned zone of the connecting partition 22.
  • the breakable part 23 may optionally include removals of material, such as holes or localized recessed areas, for example.
  • the assembly is mounted in one piece or in one piece in the turbine engine 1, then, during the first start-up of the turbine engine 1, a temperature differential is created between the arms 16 of the annular channel 6, on the one hand, and the connecting parts 18 of the support 17, on the other hand, which has the effect of breaking the breakable part 23 of the connecting partition 22 due to the stresses generated in said breakable part 23.
  • the temperature differential allowing a break in the breakable zone is for example between 200 and 500° C.
  • the breakable part 23 can be broken cold, that is to say without heating part of the assembly, before mounting the annular channel 6 and the support 17 in the turbomachine 1.
  • the breakable part 23 can be broken cold, that is to say without heating of part of the assembly, after assembly of the annular channel 6 and the support 17, in one piece, in the turbomachine 1.
  • a stress can be generated mechanically at the level of the connecting partition 22, for example by an operator, in particular by applying a shock or a sufficient force on said partition 22.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Description

DOMAINEDOMAIN

La présente invention concerne un ensemble pour une turbomachine, telle par exemple qu'un turboréacteur ou un turbopropulseur d'aéronef.The present invention relates to an assembly for a turbomachine, such as, for example, an aircraft turbojet or turboprop engine.

CONTEXTECONTEXT

La figure 1 illustre une partie d'une turbomachine 1 selon une première forme de réalisation conformément à l'art antérieur.The figure 1 illustrates part of a turbomachine 1 according to a first embodiment in accordance with the prior art.

Dans ce qui suit, les termes amont et aval sont définis par rapport au sens de circulation des gaz au sein de la turbomachine 1.In what follows, the terms upstream and downstream are defined with respect to the direction of circulation of the gases within the turbomachine 1.

La turbomachine 1 comporte une turbine amont 2 et une turbine aval 3. La turbine amont 2 est par exemple une turbine haute-pression et la turbine aval 3 est par exemple une turbine basse-pression ou une turbine libre. Chaque turbine 2, 3 comporte un rotor comprenant des pales 4. La turbomachine 1 comporte en outre un arbre 5 radialement interne, s'étendant selon l'axe A de la turbomachine 1.The turbomachine 1 comprises an upstream turbine 2 and a downstream turbine 3. The upstream turbine 2 is for example a high-pressure turbine and the downstream turbine 3 is for example a low-pressure turbine or a free turbine. Each turbine 2, 3 comprises a rotor comprising blades 4. The turbine engine 1 further comprises a radially internal shaft 5, extending along the axis A of the turbine engine 1.

La turbomachine 1 comporte de plus un canal annulaire 6 destiné à former une veine d'écoulement du flux de gaz entre deux étages de turbine 2, 3 de la turbomachine 1, ledit canal 6 étant délimité par une paroi annulaire radialement interne 7 et une paroi annulaire radialement externe 8.The turbomachine 1 further comprises an annular channel 6 intended to form a flow path for the gas flow between two turbine stages 2, 3 of the turbomachine 1, said channel 6 being delimited by a radially internal annular wall 7 and a wall radially outer ring finger 8.

Un support radialement externe 9 relie la paroi annulaire externe et un carter 10 de turbine. Le support externe 9 comporte une zone 11 souple ou élastiquement déformable, apte à autoriser le déplacement radial et/ou axial de la paroi annulaire externe 8 par rapport au carter 10.A radially outer support 9 connects the outer annular wall and a casing 10 of the turbine. The external support 9 comprises a flexible or elastically deformable zone 11, able to allow the radial and/or axial displacement of the external annular wall 8 with respect to the casing 10.

Un support radialement interne 12 s'étend radialement vers l'intérieur depuis la paroi radialement interne 7. La partie radialement interne 13 du support interne 11 entoure deux paliers 14 montés autour de l'arbre 5. Le support interne 12 comporte en outre une zone 15 souple ou élastiquement déformable, apte à autoriser le déplacement radial et/ou axial de la paroi annulaire interne 7 par rapport aux paliers 14 et à l'arbre 5.A radially internal support 12 extends radially inwards from the radially internal wall 7. The radially internal part 13 of the internal support 11 surrounds two bearings 14 mounted around the shaft 5. The internal support 12 further comprises a zone 15 flexible or elastic deformable, capable of allowing the radial and/or axial displacement of the internal annular wall 7 with respect to the bearings 14 and to the shaft 5.

L'ensemble formé par le canal annulaire 6 et les supports interne et externe 9, 12 est réalisé d'une pièce, par exemple par fonderie.The assembly formed by the annular channel 6 and the internal and external supports 9, 12 is made in one piece, for example by foundry.

En fonctionnement, les parois annulaires interne et externe 7, 8 du canal annulaire 6 sont soumises à des températures importantes, tandis que les supports interne 12 et externe 9 peuvent être soumis à des températures plus faibles. La différence de température est notamment très importante lors de la phase dite de transition, au démarrage de la turbomachine. Cette différence de température génère des dilatations différentielles entre les différentes parties du même ensemble. Les zones souples 11, 15 des supports 9, 12 permettent de compenser de telles dilatations différentielles, en autorisant un déplacement radial et/ou axial des parois annulaires interne et externe 7, 8 du canal annulaire 6 par rapport aux autres parties de l'ensemble.In operation, the internal and external annular walls 7, 8 of the annular channel 6 are subjected to high temperatures, while the internal 12 and external 9 supports can be subjected to lower temperatures. The temperature difference is in particular very large during the so-called transition phase, when the turbine engine is started. This temperature difference generates differential expansions between the different parts of the same assembly. The flexible zones 11, 15 of the supports 9, 12 make it possible to compensate for such differential expansions, by allowing radial and/or axial displacement of the internal and external annular walls 7, 8 of the annular channel 6 with respect to the other parts of the assembly. .

Cependant, une trop grande souplesse accordée aux supports 9, 12 pénaliserait le guidage de l'arbre 5, au travers des paliers 14. En effet, les supports 9, 12 assurent une fonction dite structurelle puisqu'ils ont notamment pour fonction de supporter radialement l'arbre 5, c'est-à-dire de le lier au carter 10, et d'éviter le débattement radial de l'arbre 5, en particulier sous charge.However, too much flexibility granted to the supports 9, 12 would penalize the guiding of the shaft 5, through the bearings 14. Indeed, the supports 9, 12 perform a so-called structural function since their particular function is to support radially the shaft 5, that is to say to link it to the casing 10, and to avoid the radial displacement of the shaft 5, in particular under load.

Il convient donc de trouver un compromis entre les aspects de souplesse pour autoriser les dilatations différentielles et de rigidité pour réaliser la fonction de support de l'arbre 5. Par ailleurs, les contraintes mécaniques et thermiques appliquées aux différentes parties sont importantes et pénalisent la durée de vie de l'ensemble.It is therefore necessary to find a compromise between the aspects of flexibility to allow differential expansion and rigidity to perform the support function of the shaft 5. Furthermore, the mechanical and thermal stresses applied to the different parts are significant and penalize the duration of life as a whole.

Pour pouvoir répondre aux spécifications, l'ensemble est réalisé par exemple en alliage à base de nickel de type Inconel 738, un tel matériau étant coûteux et n'étant pas réparable par rechargement de matière par soudage.In order to be able to meet the specifications, the assembly is made, for example, of a nickel-based alloy of the Inconel 738 type, such a material being expensive and not being repairable by reloading the material by welding.

Une seconde forme de réalisation connue de l'art antérieur est représentée à la figure 2. Dans cette forme de réalisation, l'ensemble comporte un canal annulaire 6 destiné à former une veine d'écoulement d'un flux de gaz entre les deux étages de turbine 2, 3 de la turbomachine 1, ledit canal 6 étant délimité par une paroi annulaire radialement interne 7 et une paroi annulaire radialement externe 8, lesdites parois 7, 8 étant reliées par des bras creux 16 s'étendant radialement.A second embodiment known from the prior art is shown in figure 2 . In this embodiment, the assembly comprises an annular channel 6 intended to form a flow path for a flow of gas between the two turbine stages 2, 3 of the turbomachine 1, said channel 6 being delimited by a wall radially inner annular wall 7 and a radially outer annular wall 8, said walls 7, 8 being connected by hollow arms 16 extending radially.

L'ensemble comporte en outre un support 17, distinct du canal annulaire 6, et comportant une partie annulaire radialement externe 9, située radialement à l'extérieur de la paroi annulaire externe 8 du canal annulaire 6, et une partie annulaire radialement interne 12, située radialement à l'intérieur de la paroi annulaire interne 7 du canal annulaire 6, les parties externe 9 et interne 12 du support 17 étant reliées par des parties de liaison 18 s'étendant radialement, chaque partie de liaison 18 traversant un bras creux 16 du canal annulaire 6.The assembly further comprises a support 17, separate from the annular channel 6, and comprising a radially outer annular part 9, located radially outside the outer annular wall 8 of the annular channel 6, and a radially inner annular part 12, located radially inside the internal annular wall 7 of the annular channel 6, the external 9 and internal 12 parts of the support 17 being connected by connecting parts 18 extending radially, each connecting part 18 passing through a hollow arm 16 of the annular canal 6.

De cette manière, il est possible de réaliser le support 17 et le canal annulaire 6 en deux matériaux différents, ce qui permet de choisir plus aisément le matériau répondant aux contraintes thermiques et mécaniques de chaque partie.In this way, it is possible to produce the support 17 and the annular channel 6 in two different materials, which makes it easier to choose the material that meets the thermal and mechanical stresses of each part.

Une telle solution reste cependant coûteuse car elle nécessite la fabrication et le montage de plusieurs pièces distinctes. En effet, une telle solution nécessite de sectoriser le support 17, l'interface et l'étanchéité entre les différents secteurs générant des contraintes supplémentaires.Such a solution, however, remains expensive because it requires the manufacture and assembly of several separate parts. Indeed, such a solution requires segmenting the support 17, the interface and the sealing between the different sectors generating additional constraints.

Les documents FR2708681A1 , US8099962B2 et US2012/227371A1 divulguent également des turbomachines selon l'art antérieur.The documents FR2708681A1 , US8099962B2 and US2012/227371A1 also disclose turbomachines according to the prior art.

RESUME DE L'INVENTIONSUMMARY OF THE INVENTION

L'invention vise à remédier à ces inconvénients, de manière simple, fiable et peu onéreuse.The invention aims to remedy these drawbacks, in a simple, reliable and inexpensive manner.

A cet effet, l'invention concerne un ensemble pour une turbomachine, comportant :

  • un canal annulaire destiné à former une veine d'écoulement d'un flux de gaz entre deux étages de turbine de la turbomachine, ledit canal étant délimité par une paroi annulaire radialement interne et une paroi annulaire radialement externe, lesdites parois étant reliées par des bras creux s'étendant radialement,
  • un support comportant une partie annulaire radialement externe, située radialement à l'extérieur de la paroi annulaire externe du canal annulaire, et une partie annulaire radialement interne, située radialement à l'intérieur de la paroi annulaire interne du canal annulaire, les parties externe et interne du support étant reliées par des parties de liaison s'étendant radialement, chaque partie de liaison traversant l'un des bras creux du canal annulaire,
    caractérisé en ce que l'une au moins des parties de liaison du support et le bras creux correspondant sont reliés l'un à l'autre par au moins une cloison de liaison, ladite cloison de liaison comportant une partie sécable apte à rompre lorsque les contraintes mécaniques dans ladite cloison de liaison sont supérieures à une valeur déterminée.
To this end, the invention relates to an assembly for a turbomachine, comprising:
  • an annular channel intended to form a flow path for a gas flow between two turbine stages of the turbomachine, said channel being delimited by a radially internal annular wall and a radially external annular wall, said walls being connected by arms radially extending hollow,
  • a support comprising a radially outer annular part, located radially outside the outer annular wall of the annular channel, and a radially inner annular part, located radially inside the inner annular wall of the annular channel, the outer and internal part of the support being connected by connecting parts extending radially, each connecting part passing through one of the hollow arms of the annular channel,
    characterized in that at least one of the connecting parts of the support and the corresponding hollow arm are connected to each other by at least one connecting partition, the said connecting partition comprising a breakable part able to break when the mechanical stresses in said connecting partition are greater than a determined value.

L'ensemble peut ainsi être réalisé en une seule pièce, par exemple par fabrication additive ou par fonderie, ce qui permet de réduire les coûts de fabrication. Après rupture de la partie sécable, le canal annulaire et le support forment deux pièces distinctes, de manière à éviter la conduction ou les ponts thermiques par contact entre lesdites pièces.The assembly can thus be produced in a single piece, for example by additive manufacturing or by foundry, which makes it possible to reduce the manufacturing costs. After breaking the breakable part, the annular channel and the support form two separate parts, so as to avoid conduction or thermal bridges by contact between said parts.

La partie sécable peut être dimensionnée pour rompre lorsque les contraintes de cisaillement dans la cloison de liaison, au niveau de la partie sécable, sont supérieures à 200 Mpa.The breakable part can be sized to break when the shear stresses in the connecting partition, at the level of the breakable part, are greater than 200 MPa.

La valeur de contrainte précitée est par exemple la valeur lorsque la cloison de liaison est à une température comprise entre 500 et 900°C, cette valeur pouvant changer avec la température.The aforementioned stress value is for example the value when the connecting partition is at a temperature between 500 and 900° C., this value being able to change with the temperature.

L'ensemble est réalisé d'une seule pièce en alliage à base de nickel, par exemple en un alliage de type C263.The assembly is made in one piece from a nickel-based alloy, for example from an alloy of the C263 type.

De préférence, l'alliage utilisé peut être rechargé par soudage. Ceci est notamment le cas d'un alliage de type C263.Preferably, the alloy used can be hardfaced by welding. This is in particular the case of an alloy of the C263 type.

La partie sécable peut être formée par une zone amincie de la cloison de liaison.The breakable part can be formed by a thinned zone of the connecting partition.

La partie sécable peut comporter des enlèvements de matière, tels par exemple que des trous ou des zones en creux localisées.The breakable part may include removals of material, such as holes or localized recessed areas, for example.

L'une au moins des parties de liaison du support peut comporter un conduit interne permettant l'amenée d'un fluide de lubrification depuis une zone située radialement à l'extérieur du canal annulaire jusque dans une zone située à l'intérieur du canal annulaire.At least one of the connection parts of the support may comprise an internal duct allowing the supply of a lubricating fluid from a zone located radially outside the annular channel to a zone located inside the annular channel .

La partie radialement interne du support peut être destinée à supporter au moins un palier. Le conduit peut ainsi permettre la lubrification dudit palier.The radially inner part of the support may be intended to support at least one bearing. The duct can thus allow the lubrication of said bearing.

Le fluide de lubrification est par exemple de la graisse ou de l'huile.The lubricating fluid is for example grease or oil.

La partie radialement interne et/ou la partie radialement externe du support peuvent comporter au moins une zone souple permettant une déformation radiale de ladite partie radialement interne ou externe.The radially internal part and/or the radially external part of the support may comprise at least one flexible zone allowing radial deformation of said radially internal or external part.

La partie radialement interne et/ou la partie radialement externe du support peuvent comporter une partie périphérie radialement fixe, reliée à chaque partie de liaison par la zone souple correspondante.The radially inner part and/or the radially outer part of the support can comprise a radially fixed peripheral part, connected to each connecting part by the corresponding flexible zone.

La zone souple peut être formée par des pattes ou des épingles élastiquement déformables.The flexible zone can be formed by elastically deformable tabs or pins.

Lesdites pattes ou épingles peuvent être orientées de façon oblique, c'est-à-dire peuvent former un angle non nul avec la direction axiale et avec la direction radiale. Ledit angle avec la direction axiale est par exemple compris entre 30 et 60°, par exemple de l'ordre de 45°.Said tabs or pins can be oriented obliquely, that is to say can form a non-zero angle with the axial direction and with the radial direction. Said angle with the axial direction is for example between 30 and 60°, for example of the order of 45°.

L'invention concerne une turbomachine, telle par exemple qu'un turboréacteur ou un turbopropulseur, comportant une turbine amont, par exemple une turbine haute-pression, et une turbine aval, par exemple une turbine basse-pression ou une turbine libre, lesdites turbines comportant chacune un rotor, la turbomachine comportant un arbre radialement interne, caractérisée en ce qu'elle comporte un ensemble du type précité, le canal annulaire formant une veine d'écoulement de gaz entre la turbine amont et la turbine aval, la partie radialement interne du support supportant au moins un palier servant au guidage de l'arbre, la partie radialement externe du support étant fixée à une partie fixe de la turbomachine, par exemple un carter de turbine.The invention relates to a turbomachine, such as for example a turbojet or a turboprop, comprising an upstream turbine, for example a high-pressure turbine, and a downstream turbine, for example a low-pressure turbine or a free turbine, said turbines each comprising a rotor, the turbomachine comprising a radially internal shaft, characterized in that it comprises an assembly of the aforementioned type, the annular channel forming a gas flow stream between the upstream turbine and the downstream turbine, the radially internal part of the support supporting at least one bearing serving to guide the shaft, the radially outer part of the support being fixed to a fixed part of the turbomachine, for example a turbine casing.

L'invention concerne également un procédé de montage et de fonctionnement d'une turbomachine du type précité, caractérisé en ce qu'il comporte les étapes consistant à :

  • monter le canal annulaire et le support dans la turbomachine,
  • effectuer un premier démarrage de la turbomachine de manière à créer un différentiel de température entre les bras du canal annulaire, d'une part, et les parties de liaison du support, d'autre part, et générer une rupture de la partie sécable de la cloison de liaison du fait des contraintes générées dans ladite partie sécable.
The invention also relates to a method for assembling and operating a turbomachine of the aforementioned type, characterized in that it comprises the steps consisting in:
  • mount the annular channel and the support in the turbomachine,
  • perform a first start-up of the turbomachine so as to create a temperature differential between the arms of the annular channel, on the one hand, and the connecting parts of the support, on the other hand, and generate a break in the breakable part of the connecting partition due to the stresses generated in said breakable part.

Le différentiel de température permettant une rupture de la zone sécable est par exemple compris entre 200 et 500 °C.The temperature differential allowing a break in the breakable zone is for example between 200 and 500° C.

En variante, la partie sécable peut être rompue à froid, c'est-à-dire sans échauffement d'une partie de l'ensemble, avant montage du canal annulaire et du support dans la turbomachine.As a variant, the breakable part can be broken cold, that is to say without heating part of the assembly, before mounting the annular channel and the support in the turbomachine.

Selon une autre variante, la partie sécable peut être rompue à froid, c'est-à-dire sans échauffement d'une partie de l'ensemble, après montage du canal annulaire et du support dans la turbomachine.According to another variant, the breakable part can be broken cold, that is to say without heating part of the assembly, after mounting the annular channel and the support in the turbomachine.

Pour cela, une contrainte peut être générée mécaniquement au niveau de la cloison de liaison, par exemple par un opérateur, notamment par application d'un choc ou d'un effort suffisant sur ladite cloison.For this, a stress can be generated mechanically at the level of the connecting partition, for example by an operator, in particular by applying a shock or a sufficient force to said partition.

L'invention sera mieux comprise et d'autres détails, caractéristiques et avantages de l'invention apparaîtront à la lecture de la description suivante faite à titre d'exemple non limitatif en référence aux dessins annexés.The invention will be better understood and other details, characteristics and advantages of the invention will appear on reading the following description given by way of non-limiting example with reference to the appended drawings.

BREVE DESCRIPTION DES FIGURESBRIEF DESCRIPTION OF FIGURES

  • la figure 1 est demie-vue schématique en coupe axiale d'une partie d'une turbine d'une turbomachine selon une première forme de réalisation de l'art antérieur ;the figure 1 is a diagrammatic half-view in axial section of part of a turbine of a turbomachine according to a first embodiment of the prior art;
  • la figure 2 est une vue correspondant à la figure 1, illustrant une seconde forme de réalisation de l'art antérieur ;the figure 2 is a view corresponding to the figure 1 , illustrating a second embodiment of the prior art;
  • la figure 3 est une vue correspondant à la figure 1, illustrant une forme de réalisation de l'invention ;the picture 3 is a view corresponding to the figure 1 , illustrating an embodiment of the invention;
  • la figure 4 est vue en perspective et en coupe axiale, d'une partie d'un ensemble selon l'invention ;the figure 4 is seen in perspective and in axial section, of part of an assembly according to the invention;
  • la figure 5 est une vue en perspective d'une partie de l'ensemble de la figure 4, certains éléments ayant été retirés afin d'améliorer la visibilité des éléments représentés ;the figure 5 is a perspective view of part of the whole figure 4 , some elements having been removed in order to improve the visibility of the elements represented;
  • la figure 6 est une vue en perspective d'une partie de la cloison de liaison.the figure 6 is a perspective view of part of the connecting bulkhead.
DESCRIPTION DETAILLEEDETAILED DESCRIPTION

La figure 3 représente une partie d'une turbomachine 1 selon une forme de réalisation de l'invention. Celle-ci comporte une turbine amont 2 et une turbine aval 3. La turbine amont 2 est par exemple une turbine haute-pression et la turbine aval 3 est par exemple une turbine basse-pression ou une turbine libre. Chaque turbine 2, 3 comporte un rotor comportant des pales 4. La turbomachine 1 comporte également un arbre 5 radialement interne, s'étendant selon l'axe A de la turbomachine.The picture 3 represents a part of a turbomachine 1 according to one embodiment of the invention. This comprises an upstream turbine 2 and a downstream turbine 3. The upstream turbine 2 is for example a high-pressure turbine and the downstream turbine 3 is for example a low-pressure turbine or a free turbine. Each turbine 2, 3 comprises a rotor comprising blades 4. The turbomachine 1 also comprises a radially internal shaft 5, extending along the axis A of the turbomachine.

La turbomachine 1 comporte de plus un ensemble comprenant un canal annulaire 6 destiné à former une veine d'écoulement d'un flux de gaz entre les deux étages 2, 3 de turbine de la turbomachine 1, ledit canal 6 étant délimité par une paroi annulaire radialement interne 7 et une paroi annulaire radialement externe 8, lesdites parois 7, 8 étant reliées par des bras creux 16 s'étendant radialement.The turbomachine 1 further comprises an assembly comprising an annular channel 6 intended to form a flow path for a flow of gas between the two stages 2, 3 of the turbine of the turbomachine 1, said channel 6 being delimited by an annular wall radially inner 7 and a radially outer annular wall 8, said walls 7, 8 being connected by hollow arms 16 extending radially.

L'ensemble, également visible à la figure 4, comporte en outre un support 17 comportant une partie annulaire radialement externe 9, située radialement à l'extérieur de la paroi annulaire externe 8 du canal annulaire 6, et une partie annulaire radialement interne 12, située radialement à l'intérieur de la paroi annulaire interne 7 du canal annulaire 6, les parties externe et interne 9, 12 du support 17 étant reliées par des parties de liaison 18 s'étendant radialement, chaque partie de liaison 18 traversant l'un des bras creux 16 du canal annulaire 6. Les bras creux 16 et les parties de liaison 18 sont régulièrement répartis sur la périphérie.The whole, also visible at the figure 4 , further comprises a support 17 comprising a radially outer annular part 9, located radially outside the outer annular wall 8 of the annular channel 6, and a radially inner annular part 12, located radially inside the annular wall 7 of the annular channel 6, the outer and inner parts 9, 12 of the support 17 being connected by connecting parts 18 extending radially, each connecting part 18 passing through one of the hollow arms 16 of the annular channel 6. hollow arm 16 and the connecting parts 18 are regularly distributed over the periphery.

La partie radialement interne 12 et la partie radialement externe 9 du support 17 comportent chacune une zone souple 11, 15 permettant une déformation radiale de ladite partie radialement interne ou externe 12, 9.The radially internal part 12 and the radially external part 9 of the support 17 each comprise a flexible zone 11, 15 allowing radial deformation of said radially internal or external part 12, 9.

La partie radialement interne 12 comporte une bride annulaire 19 radialement externe, s'étendant radialement, et fixée au carter 10 par l'intermédiaire de vis ou de rivets par exemple. Ladite bride 19 est reliée à chaque partie de liaison 18 par la zone souple 11 correspondante. Cette zone souple 11 est formée par des pattes ou par des épingles 20 élastiquement déformables.The radially inner part 12 comprises a radially outer annular flange 19, extending radially, and fixed to the casing 10 by means of screws or rivets for example. Said flange 19 is connected to each connecting part 18 by the corresponding flexible zone 11. This flexible zone 11 is formed by lugs or by elastically deformable pins 20.

Lesdites pattes ou épingles 20 peuvent être orientées de façon oblique, c'est-à-dire peuvent former un angle non nul avec la direction axiale et avec la direction radiale. Ledit angle avec la direction axiale est par exemple compris entre 30 et 60°, par exemple de l'ordre de 45°.Said tabs or pins 20 can be oriented obliquely, that is to say can form a non-zero angle with the axial direction and with the radial direction. Said angle with the axial direction is for example between 30 and 60°, for example of the order of 45°.

Par ailleurs, la partie radialement interne 12 du support 17 comporte des parties annulaires 13a, 13b s'étendant axialement, destinées à entourer chacune l'un des paliers 14. Chaque partie annulaire 13a, 13b est reliée aux parties de liaison 18 par des zones souples 15a, 15b obliques ou tronconiques. Chaque zone souple oblique ou tronconique 15a, 15b forme un angle non nul avec les directions axiale et radiale.Furthermore, the radially internal part 12 of the support 17 comprises annular parts 13a, 13b extending axially, each intended to surround one of the bearings 14. Each annular part 13a, 13b is connected to the connecting parts 18 by zones flexible 15a, 15b oblique or tapered. Each flexible oblique or tapered zone 15a, 15b forms a non-zero angle with the axial and radial directions.

L'une au moins des parties de liaison 18 du support 17 comporte un conduit interne 21 permettant l'amenée d'un fluide de lubrification depuis une zone située radialement à l'extérieur du canal annulaire 6 jusque dans une zone située en regard des paliers 14. Le fluide de lubrification est par exemple de la graisse ou de l'huile.At least one of the connecting parts 18 of the support 17 comprises an internal duct 21 allowing the supply of a lubricating fluid from a zone located radially outside the annular channel 6 into a area located opposite the bearings 14. The lubricating fluid is for example grease or oil.

Chaque partie de liaison peut présenter deux parties rectilignes 18a, 18b formant un angle l'une par rapport à l'autre. Bien entendu, d'autres formes sont également possibles.Each connecting part can have two rectilinear parts 18a, 18b forming an angle with respect to each other. Of course, other shapes are also possible.

Comme cela est mieux visible à la figure 5, l'une au moins des parties de liaison 18 et le bras creux 16 correspondant sont reliés l'un à l'autre par au moins une cloison de liaison 22, ladite cloison de liaison 22 comportant une partie sécable 23 apte à rompre lorsque les contraintes mécaniques dans ladite cloison de liaison 22 sont supérieures à une valeur déterminée.As is best seen at the figure 5 , at least one of the connecting parts 18 and the corresponding hollow arm 16 are connected to each other by at least one connecting partition 22, said connecting partition 22 comprising a breakable part 23 capable of breaking when the mechanical stresses in said connecting partition 22 are greater than a determined value.

On notera, que, mis à part par l'intermédiaire de la cloison de liaison 22, la partie de liaison 18 n'est pas en contact avec la surface du bras de liaison 16, de façon à limiter les échanges thermiques.It will be noted that, apart from via the connecting partition 22, the connecting part 18 is not in contact with the surface of the connecting arm 16, so as to limit heat exchange.

La partie sécable 23 peut être dimensionnée pour rompre lorsque les contraintes de cisaillement dans la cloison de liaison 22, au niveau de la partie sécable 23, sont supérieures à 200 Mpa. Cette valeur peut changer avec la température et peut par exemple être définie à une température comprise entre 500°C et 900°C.The breakable part 23 can be sized to break when the shear stresses in the connecting partition 22, at the level of the breakable part 23, are greater than 200 MPa. This value can change with temperature and can for example be set at a temperature between 500°C and 900°C.

L'ensemble formé par le canal 6 et le support 17 peut ainsi être réalisé en une seule pièce, par exemple par fabrication additive ou par fonderie, ce qui permet de réduire les coûts de fabrication. Après rupture de la partie sécable 23, le canal annulaire 6 et le support 17 forment deux pièces distinctes, de manière à éviter la conduction ou les ponts thermiques par contact entre lesdites pièces 6, 17.The assembly formed by the channel 6 and the support 17 can thus be produced in a single piece, for example by additive manufacturing or by foundry, which makes it possible to reduce the manufacturing costs. After breakage of the breakable part 23, the annular channel 6 and the support 17 form two separate parts, so as to avoid conduction or thermal bridges by contact between said parts 6, 17.

L'ensemble est réalisé d'une seule pièce en alliage à base de nickel, par exemple en un alliage de type C263.The assembly is made in one piece from a nickel-based alloy, for example from an alloy of the C263 type.

Comme cela est mieux visible à la figure 6, la partie sécable 23 de la cloison de liaison 22 est formée par une zone amincie de la cloison de liaison 22.As is best seen at the figure 6 , the breakable part 23 of the connecting partition 22 is formed by a thinned zone of the connecting partition 22.

La partie sécable 23 peut éventuellement comporter des enlèvements de matière, tels par exemple que des trous ou des zones en creux localisées.The breakable part 23 may optionally include removals of material, such as holes or localized recessed areas, for example.

Selon une première forme de réalisation, l'ensemble est monté d'un seul tenant ou d'un seul bloc dans la turbomachine 1, puis, lors du premier démarrage de la turbomachine 1, un différentiel de température se créé entre les bras 16 du canal annulaire 6, d'une part, et les parties de liaison 18 du support 17, d'autre part, ce qui a pour effet de rompre la partie sécable 23 de la cloison de liaison 22 du fait des contraintes générées dans ladite partie sécable 23.According to a first embodiment, the assembly is mounted in one piece or in one piece in the turbine engine 1, then, during the first start-up of the turbine engine 1, a temperature differential is created between the arms 16 of the annular channel 6, on the one hand, and the connecting parts 18 of the support 17, on the other hand, which has the effect of breaking the breakable part 23 of the connecting partition 22 due to the stresses generated in said breakable part 23.

Le différentiel de température permettant une rupture de la zone sécable est par exemple compris entre 200 et 500 °C.The temperature differential allowing a break in the breakable zone is for example between 200 and 500° C.

En variante, la partie sécable 23 peut être rompue à froid, c'est-à-dire sans échauffement d'une partie de l'ensemble, avant montage du canal annulaire 6 et du support 17 dans la turbomachine 1.As a variant, the breakable part 23 can be broken cold, that is to say without heating part of the assembly, before mounting the annular channel 6 and the support 17 in the turbomachine 1.

Selon une autre variante, la partie sécable 23 peut être rompue à froid, c'est-à-dire sans échauffement d'une partie de l'ensemble, après montage du canal annulaire 6 et du support 17, d'un seul tenant, dans la turbomachine 1.According to another variant, the breakable part 23 can be broken cold, that is to say without heating of part of the assembly, after assembly of the annular channel 6 and the support 17, in one piece, in the turbomachine 1.

Pour cela, une contrainte peut être générée mécaniquement au niveau de la cloison de liaison 22, par exemple par un opérateur, notamment par application d'un choc ou d'un effort suffisant sur ladite cloison 22.For this, a stress can be generated mechanically at the level of the connecting partition 22, for example by an operator, in particular by applying a shock or a sufficient force on said partition 22.

Claims (10)

  1. Assembly for a turbomachine (1), comprising:
    - an annular channel (6) intended to form a flow path for a gas stream between two turbine stages (2, 3) of the turbomachine (1), said channel (6) being delimited by a radially inner annular wall (7) and a radially outer annular wall (8), said walls (7, 8) being connected by radially extending hollow arms (16),
    - a support (17) having a radially outer annular portion (9), located radially outside the outer annular wall (8) of the annular channel (6), and a radially inner annular portion (12), located radially inside the inner annular wall (7) of the annular channel (6), the outer (9) and inner (12) portions of the support (17) being connected by radially extending connecting portions (18), each connecting portion (18) passing through a hollow arm (16) of the annular channel (6),
    characterised in that at least one of the connecting parts (18) of the support (17) and the corresponding hollow arm (16) are connected to each other by at least one connecting partition (22), said connecting partition (22) having a breakable part (23) capable of breaking when the mechanical stresses in said connecting partition (22) are greater than a predetermined value.
  2. Assembly according to claim 1, characterised in that the breakable portion (23) is dimensioned to break when the shear stresses in the connecting partition (22) at the breakable portion (23) are greater than 200 MPa.
  3. Assembly according to claim 1 or 2, characterised in that the assembly is made in one piece from a nickel-based alloy, for example an alloy of type C263.
  4. Assembly according to one of claims 1 to 3, characterised in that the breakable portion (23) is formed by at least one thinned part of the connecting partition (22) .
  5. Assembly according to one of claims 1 to 4, characterised in that the breakable part (23) has material removals such as, for example, holes or localized depressed areas.
  6. Assembly according to one of claims 1 to 5, characterised in that at least one of the connecting parts (18) of the support (17) may have an internal conduit (21) for the supply of a lubricating fluid from an area located radially outside the annular channel (6) up to an area located inside the annular channel (6).
  7. Assembly according to one of claims 1 to 6, characterised in that the radially inner part (12) and/or the radially outer part (9) of the support (17) comprise at least one flexible zone (15, 11) allowing radial deformation of said radially inner or outer part (9, 12).
  8. Assembly according to claim 7, characterised in that the radially inner part (12) and/or the radially outer part (9) of the support (17) has a radially fixed peripheral part (19, 13a, 13b), connected to each connecting part (18) by the corresponding flexible zone (11, 15).
  9. Turbomachine (1), such as for example a turbojet or turboprop, comprising an upstream turbine (2), for example a high-pressure turbine, and a downstream turbine (3), for example a low-pressure turbine or a free turbine, said turbines (2, 3) each comprising a rotor, the turbomachine (1) comprising a radially inner shaft (5), characterised in that it comprises an assembly according to one of claims 1 to 8, the annular channel (6) forming a gas flow path between the upstream turbine (2) and the downstream turbine (3), the radially inner part (12) of the support (17) supporting at least one bearing (14) serving to guide the shaft (5), the radially outer part (9) of the support (17) being fixed to a fixed part (10) of the turbomachine (1), for example a turbine casing (10).
  10. Method of assembling and operating a turbomachine
    (1) according to claim 9, characterised in that it comprises the following steps:
    - mounting the annular channel (6) and the support (17) in the turbomachine (1),
    - performing a first start-up of the turbomachine (1) so as to create a temperature differential between the arms (16) of the annular channel (6), on the one hand, and the connecting parts (18) of the support (17), on the other hand, and to generate a break in the breakable part (23) of the connecting partition (22) due to the stresses generated in said breakable part (23).
EP19717517.7A 2018-02-28 2019-02-28 Assembly for a turbomachine Active EP3759319B1 (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
PL19717517T PL3759319T3 (en) 2018-02-28 2019-02-28 Assembly for a turbomachine

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR1851776A FR3078370B1 (en) 2018-02-28 2018-02-28 SET FOR A TURBOMACHINE
PCT/FR2019/050462 WO2019166742A1 (en) 2018-02-28 2019-02-28 Assembly for a turbomachine

Publications (2)

Publication Number Publication Date
EP3759319A1 EP3759319A1 (en) 2021-01-06
EP3759319B1 true EP3759319B1 (en) 2022-01-12

Family

ID=62816675

Family Applications (1)

Application Number Title Priority Date Filing Date
EP19717517.7A Active EP3759319B1 (en) 2018-02-28 2019-02-28 Assembly for a turbomachine

Country Status (7)

Country Link
US (1) US11181009B2 (en)
EP (1) EP3759319B1 (en)
CN (1) CN111801487B (en)
CA (1) CA3091499A1 (en)
FR (1) FR3078370B1 (en)
PL (1) PL3759319T3 (en)
WO (1) WO2019166742A1 (en)

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3120900B1 (en) 2021-03-18 2023-02-10 Safran Aircraft Engines CENTERING AND GUIDE DEVICE FOR AN AIRCRAFT TURBOMACHINE SHAFT
FR3120904B1 (en) 2021-03-18 2023-03-24 Safran Aircraft Engines CENTERING AND GUIDE DEVICE FOR AN AIRCRAFT TURBOMACHINE SHAFT
FR3120899B1 (en) 2021-03-18 2023-05-26 Safran Aircraft Engines CENTERING AND GUIDE DEVICE FOR AN AIRCRAFT TURBOMACHINE SHAFT
FR3120902B1 (en) 2021-03-18 2023-03-10 Safran Aircraft Engines CENTERING AND GUIDE DEVICE FOR AN AIRCRAFT TURBOMACHINE SHAFT

Family Cites Families (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5451116A (en) * 1992-06-09 1995-09-19 General Electric Company Tripod plate for turbine flowpath
GB2326679B (en) * 1997-06-25 2000-07-26 Rolls Royce Plc Ducted fan gas turbine engine
US6240719B1 (en) * 1998-12-09 2001-06-05 General Electric Company Fan decoupler system for a gas turbine engine
GB2360069B (en) * 2000-03-11 2003-11-26 Rolls Royce Plc Ducted fan gas turbine engine
US6402469B1 (en) * 2000-10-20 2002-06-11 General Electric Company Fan decoupling fuse
GB2444935B (en) * 2006-12-06 2009-06-10 Rolls Royce Plc A turbofan gas turbine engine
US8099962B2 (en) * 2008-11-28 2012-01-24 Pratt & Whitney Canada Corp. Mid turbine frame system and radial locator for radially centering a bearing for gas turbine engine
US8979477B2 (en) * 2011-03-09 2015-03-17 General Electric Company System for cooling and purging exhaust section of gas turbine engine
US9777596B2 (en) * 2013-12-23 2017-10-03 Pratt & Whitney Canada Corp. Double frangible bearing support

Also Published As

Publication number Publication date
CN111801487B (en) 2022-06-28
FR3078370A1 (en) 2019-08-30
FR3078370B1 (en) 2020-02-14
US11181009B2 (en) 2021-11-23
WO2019166742A1 (en) 2019-09-06
US20200408109A1 (en) 2020-12-31
CA3091499A1 (en) 2019-09-06
PL3759319T3 (en) 2022-03-21
EP3759319A1 (en) 2021-01-06
CN111801487A (en) 2020-10-20

Similar Documents

Publication Publication Date Title
EP3759319B1 (en) Assembly for a turbomachine
EP1840339B1 (en) Device for attaching ring sectors around a turbine wheel of a turbomachine
CA2889751C (en) Exhaust housing hub for a turbomachine
EP2917519B1 (en) Air exhaust tube holder in a turbomachine
EP3775501B1 (en) Cooling device for a turbine of a turbomachine
EP2085574A1 (en) Centring of a part inside a shaft
CA2931044C (en) Device for centering and guiding in rotation with a turbomachine shaft and comprising methods for retaining the bearing outer race
EP3256698B1 (en) Retainer nut for axial blockage of a bearing in a turbomachine
EP3610136B1 (en) Device for centring and rotationally guiding a turbomachine shaft comprising means for an axial retention of outer ring of a bearing
FR3071546B1 (en) AXIAL RETENTION OF THE BLOWER TREE IN A GAS TURBINE ENGINE
FR3077097A1 (en) COOLING DEVICE FOR TURBINE OF A TURBOMACHINE
EP2917518B1 (en) Air exhaust tube holder in a turbomachine
EP2071141B1 (en) Watertigh fixing of a bearing support in a turbomachine
FR2971022A1 (en) COMPRESSOR RECTIFIER STAGE FOR A TURBOMACHINE
EP2071130B1 (en) Installation of tubes for pressurising an internal enclosure in a turbomachine
EP3824221B1 (en) Assembly for a turbomachine
EP4090833B1 (en) Assembly for a turbine engine
FR2903151A1 (en) Exhaust case ventilation device for e.g. aircraft`s jet engine, has slits delimiting blades that are supported on stiffener and freely deformable in radial direction under effect of differential thermal dilatation between case and trim
FR3055354B1 (en) TURBOMACHINE COMPRISING MEANS FOR SEALING AND METHOD FOR MOUNTING THE CORRESPONDING TURBOMACHINE
FR3065481A1 (en) TURBINE ASSEMBLY, IN PARTICULAR FOR A TURBOMACHINE
FR3094031A1 (en) SET FOR A TURBOMACHINE
EP3853445B1 (en) Turbine seal
FR3081924A1 (en) TURBOMACHINE FOR AN AIRCRAFT COMPRISING A PRESSURIZED FLUID CONDUIT SURROUNDED BY A BRAIDED OR WOVEN METAL SHEATH
EP3880940B1 (en) Device for centring and guiding in rotation a rotating part with interlaced arms
FR3055310A1 (en) TURBOMACHINE WITH OIL GUIDING DEVICE AND METHOD OF DISASSEMBLING THE TURBOMACHINE

Legal Events

Date Code Title Description
STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: UNKNOWN

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

Free format text: STATUS: THE INTERNATIONAL PUBLICATION HAS BEEN MADE

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

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

Free format text: STATUS: REQUEST FOR EXAMINATION WAS MADE

17P Request for examination filed

Effective date: 20200825

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

AX Request for extension of the european patent

Extension state: BA ME

DAV Request for validation of the european patent (deleted)
DAX Request for extension of the european patent (deleted)
GRAP Despatch of communication of intention to grant a patent

Free format text: ORIGINAL CODE: EPIDOSNIGR1

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

Free format text: STATUS: GRANT OF PATENT IS INTENDED

INTG Intention to grant announced

Effective date: 20210930

GRAS Grant fee paid

Free format text: ORIGINAL CODE: EPIDOSNIGR3

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

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

Free format text: STATUS: THE PATENT HAS BEEN GRANTED

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

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: CH

Ref legal event code: EP

REG Reference to a national code

Ref country code: DE

Ref legal event code: R096

Ref document number: 602019010900

Country of ref document: DE

REG Reference to a national code

Ref country code: IE

Ref legal event code: FG4D

Free format text: LANGUAGE OF EP DOCUMENT: FRENCH

REG Reference to a national code

Ref country code: AT

Ref legal event code: REF

Ref document number: 1462517

Country of ref document: AT

Kind code of ref document: T

Effective date: 20220215

REG Reference to a national code

Ref country code: LT

Ref legal event code: MG9D

REG Reference to a national code

Ref country code: NL

Ref legal event code: MP

Effective date: 20220112

REG Reference to a national code

Ref country code: AT

Ref legal event code: MK05

Ref document number: 1462517

Country of ref document: AT

Kind code of ref document: T

Effective date: 20220112

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

Ref country code: NL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

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

Ref country code: SE

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: RS

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: PT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220512

Ref country code: NO

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220412

Ref country code: LT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: HR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: ES

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: BG

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220412

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

Ref country code: LV

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: GR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220413

Ref country code: FI

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: AT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

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

Ref country code: IS

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220512

REG Reference to a national code

Ref country code: DE

Ref legal event code: R097

Ref document number: 602019010900

Country of ref document: DE

REG Reference to a national code

Ref country code: CH

Ref legal event code: PL

REG Reference to a national code

Ref country code: BE

Ref legal event code: MM

Effective date: 20220228

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

Ref country code: SM

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: SK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: RO

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: MC

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: LU

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

Effective date: 20220228

Ref country code: EE

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: DK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

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

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

Ref country code: AL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

26N No opposition filed

Effective date: 20221013

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

Ref country code: LI

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

Effective date: 20220228

Ref country code: IE

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

Effective date: 20220228

Ref country code: CH

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

Effective date: 20220228

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

Ref country code: SI

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: BE

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

Effective date: 20220228

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

Ref country code: MK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

Ref country code: CY

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112

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

Ref country code: DE

Payment date: 20240123

Year of fee payment: 6

Ref country code: CZ

Payment date: 20240125

Year of fee payment: 6

Ref country code: GB

Payment date: 20240123

Year of fee payment: 6

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

Ref country code: HU

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT; INVALID AB INITIO

Effective date: 20190228

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

Ref country code: PL

Payment date: 20240126

Year of fee payment: 6

Ref country code: IT

Payment date: 20240123

Year of fee payment: 6

Ref country code: FR

Payment date: 20240123

Year of fee payment: 6

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

Ref country code: TR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20220112