EP3759319B1 - Ensemble pour une turbomachine - Google Patents

Ensemble pour une turbomachine Download PDF

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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)
English (en)
French (fr)
Other versions
EP3759319A1 (fr
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/pl
Publication of EP3759319A1 publication Critical patent/EP3759319A1/fr
Application granted granted Critical
Publication of EP3759319B1 publication Critical patent/EP3759319B1/fr
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Anticipated expiration legal-status Critical

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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)
EP19717517.7A 2018-02-28 2019-02-28 Ensemble pour une turbomachine Active EP3759319B1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
PL19717517T PL3759319T3 (pl) 2018-02-28 2019-02-28 Zespół do maszyny wirowej

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR1851776A FR3078370B1 (fr) 2018-02-28 2018-02-28 Ensemble pour une turbomachine
PCT/FR2019/050462 WO2019166742A1 (fr) 2018-02-28 2019-02-28 Ensemble pour une turbomachine

Publications (2)

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

Family

ID=62816675

Family Applications (1)

Application Number Title Priority Date Filing Date
EP19717517.7A Active EP3759319B1 (fr) 2018-02-28 2019-02-28 Ensemble pour une turbomachine

Country Status (7)

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

Families Citing this family (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3120902B1 (fr) 2021-03-18 2023-03-10 Safran Aircraft Engines Dispositif de centrage et de guidage d’un arbre de turbomachine d’aeronef
FR3120900B1 (fr) 2021-03-18 2023-02-10 Safran Aircraft Engines Dispositif de centrage et de guidage d’un arbre de turbomachine d’aeronef
FR3120899B1 (fr) 2021-03-18 2023-05-26 Safran Aircraft Engines Dispositif de centrage et de guidage d’un arbre de turbomachine d’aeronef
FR3120904B1 (fr) 2021-03-18 2023-03-24 Safran Aircraft Engines Dispositif de centrage et de guidage d’un arbre de turbomachine d’aeronef
GB202307284D0 (en) * 2023-05-16 2023-06-28 Rolls Royce Plc A single-piece annular vane for a gas turbine engine

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

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