Safran Aircraft Engines

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IPC Class
F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings 275
F01D 5/14 - Form or construction 262
F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector 241
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion 198
F01D 5/30 - Fixing blades to rotors; Blade roots 161
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1.

ASSEMBLY FOR A TURBOMACHINE BLADE, COMPRISING A FASTENER, DEFINING A POCKET, AND A SHIM ADAPTED TO BE RECEIVED IN THE POCKET AT THE SAME TIME AS A ROOT OF THE BLADE

      
Application Number 17769617
Status Pending
Filing Date 2020-10-16
First Publication Date 2024-07-11
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Courtier, Vivien Mickaël
  • Jacquemard, Christophe Paul
  • Joudon, Vincent

Abstract

Assembly for a turbomachine blade, comprising: a fastener (9) defining a pocket (10) for receiving a blade root (11), and comprising two flanks defining between them a passage leading into the pocket (10) and forming stops preventing the root (11) from exiting from the pocket (10) via the passage, a shim (26a, 26b) adapted to be received in the pocket (10) with the root (11), the shim (26a, 26b) having a first support surface (28a, 28b) to bear on the root (11) and a second support surface (30a, 30b) opposite the first support surface (28a, 28b) to bear against the fastener (9), the support surfaces being fixed relative to each other and oriented such that movement of the shim (26a, 26b) in the pocket (10) relative to the fastener (9) varies the value of force exerted by the shim (26a, 26b) on the root (11).

IPC Classes  ?

  • F01D 5/30 - Fixing blades to rotors; Blade roots
  • B64C 11/06 - Blade mountings for variable-pitch blades
  • F01D 5/32 - Locking, e.g. by final locking-blades or keys
  • F01D 7/00 - Rotors with blades adjustable in operation; Control thereof

2.

GUIDE VANES ASSEMBLY WITH POSITION-MAINTAINING DEVICE

      
Application Number 18546169
Status Pending
Filing Date 2022-02-08
First Publication Date 2024-07-11
Owner
  • SAFRAN AIRCRAFT ENGINES (France)
  • SAFRAN CERAMICS (France)
Inventor
  • Congratel, Sébastien Serge Francis
  • Gaillard, Aurélien
  • Gimat, Matthieu Arnaud
  • Jarrossay, Clément
  • Le Cair, David René Pierre

Abstract

A turbine engine nozzle extending about an axis and including a vane assembly mounted between an outer shroud and an inner shroud; and a position retaining device configured to be placed in an elastically prestressed state between the inner shroud and an inner platform of the vane assembly to apply a stress on the vane assembly towards the outer shroud. The contact between the inner shroud and the inner platform is obtained by one same surface of the position retaining device.

IPC Classes  ?

  • F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector

3.

TURBOMACHINE DISTRIBUTOR COMPRISING A GAS REINTRODUCTION DUCT WITH A TANGENTIAL COMPONENT

      
Application Number 18546534
Status Pending
Filing Date 2022-02-14
First Publication Date 2024-07-11
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Lemonnier, Jérôme Claude George
  • Boisnault, Franck Davy
  • Van Noort, Antoine Bruno

Abstract

A distributor for a turbine of a turbomachine. The distributor includes a radially inner platform, a blade and a root which is coated with a track made of an abradable material. According to the invention, the distributor includes a leakage gas reintroduction duct. The duct includes an inlet through the track made of an abradable material, an outlet opening through a downstream surface of the root, and a duct intermediate portion which extends from the inlet up to the outlet. The duct intermediate portion is orientated with a tangential component.

IPC Classes  ?

  • F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
  • F01D 11/02 - Preventing or minimising internal leakage of working fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type

4.

DISTRIBUTED PROPULSION WITH OFFSET PROPULSION MODULES

      
Application Number 17769554
Status Pending
Filing Date 2020-10-14
First Publication Date 2024-07-11
Owner
  • INSTITUT SUPERIEUR DE L'AERONAUTIQUE ET DE L'ESPAC (France)
  • SAFRAN (France)
  • SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Gonidec, Patrick
  • Aguirre, Miguel Angel
  • Duplaa, Sébastien
  • Robic, Bernard
  • Rodriguez, Benoit
  • Tantot, Nicolas Jérôme Jean

Abstract

A propulsion unit including: a gas generator including a compressor, a combustion chamber, a turbine, at least two compressed-air propulsion modules, each propulsion module including: a fan, a compressed-air turbine configured to drive the rotation of the fan, a manifold allowing the respective turbines of the compressed-air propulsion modules to be supplied with compressed air, wherein the manifold is configured to collect and mix: at least a portion of the flow that has passed through the combustion chamber of the gas generator, typically as it leaves the turbine, and at least one bypass flow, the bypass flow being a flow of air which is not passed through the combustion chamber of the gas generator.

IPC Classes  ?

  • F02C 3/04 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
  • B64D 27/10 - Aircraft characterised by the type or position of power plant of gas-turbine type
  • F04D 17/02 - Radial-flow pumps specially adapted for elastic fluids, e.g. centrifugal pumps; Helico-centrifugal pumps specially adapted for elastic fluids having non-centrifugal stages, e.g. centripetal

5.

TURBOJET ENGINE NOZZLE RING FOR AN AIRCRAFT

      
Application Number 18278991
Status Pending
Filing Date 2022-02-18
First Publication Date 2024-07-11
Owner Safran Aircraft Engines (France)
Inventor
  • Clement, Alice Marie
  • Grandin, Paul
  • Marchand, Clément
  • Patel, Vijeay
  • Sanchez Gomez, Mireya

Abstract

The turbomachine subassembly comprises:—at least one tow-pressure nozzle ring, the nozzle ring comprising a root (38) integral with the nozzle ring;—a clamp (40) configured so as to fasten the root to a casing of the turbomachine; and—at least two pads (150) interposed between the root and the clamp, each pad being assembled in a form-fitting manner with one of either the root or the clamp.

IPC Classes  ?

  • F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
  • F01D 25/00 - Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
  • F01D 25/28 - Supporting or mounting arrangements, e.g. for turbine casing
  • F02K 1/52 - Nozzles specially constructed for positioning adjacent to another nozzle or to a fixed member, e.g. fairing

6.

ACCESSORY GEARBOX FOR AN AIRCRAFT TURBINE ENGINE COMPRISING A VERY LOW-SPEED ENGINE AND USE METHOD

      
Application Number 18546903
Status Pending
Filing Date 2022-02-24
First Publication Date 2024-07-11
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor Dautreppe, Frédéric

Abstract

An accessory gearbox for an aircraft turbine engine having a radial shaft connecting a high-pressure body of the turbine engine to a gear train having at least one mixed shaft coupled to a high-speed device. The accessory gearbox has a clutch system having a low-speed motor mounted thereon and configured to mate with the mixed shaft such that: in an engaged position of the clutch, the low-speed motor is coupled to the mixed shaft so as to drive the high-pressure body at low speed when the turbine engine is stationary and, in a disengaged position of the clutch, the low-speed motor is decoupled from the mixed shaft so as to be protected when the turbine engine is in operation.

IPC Classes  ?

  • F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
  • B64D 35/00 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions
  • F01D 15/12 - Combinations with mechanical gearing

7.

METHOD AND DEVICE FOR GENERATING AN ALARM FOR MONITORING AN OIL FILTER IN AN AIRCRAFT

      
Application Number 18547642
Status Pending
Filing Date 2022-02-18
First Publication Date 2024-07-11
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Muller, Jean-Louis
  • Mallet, Thibault Maxime Adrien

Abstract

A method for generating an alarm is implemented by an alarm-generating device installed in an aircraft. It comprises: a step of obtaining information delivered by a sensor of said device and representative of an oil flow rate inside an oil filter of the aircraft; a step of obtaining a temperature delivered by a sensor of said device and representative of a temperature of the oil inside said filter; a step of obtaining a differential pressure at said filter; and a step of generating an alarm if the differential pressure exceeds an alarm threshold obtained on the basis of said information and of said temperature.

IPC Classes  ?

  • B01D 35/143 - Filter condition indicators
  • B64D 45/00 - Aircraft indicators or protectors not otherwise provided for

8.

METHOD FOR MANUFACTURING A FAN DISK WITH A PART BY ADDITIVE MANUFACTURING

      
Application Number 18550395
Status Pending
Filing Date 2022-03-11
First Publication Date 2024-07-11
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Hasquenoph, Alexandre Michel Henri
  • Guessine, Hamza
  • Smith, Jeremy
  • Vigourt, Antonin

Abstract

A method for manufacturing a fan disk of a turbine engine includes forging a first cylindrical shape of axis X, machining the first cylindrical shape to obtain a second cylindrical shape, adding by additive manufacturing at least one relief clement to the second cylindrical shape.

IPC Classes  ?

  • B22F 3/17 - Manufacture of workpieces or articles from metallic powder characterised by the manner of compacting or sintering; Apparatus specially adapted therefor by forging
  • B22F 5/00 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
  • B33Y 10/00 - Processes of additive manufacturing

9.

PLATFORM FOR A FAN ASSEMBLY

      
Application Number 18561861
Status Pending
Filing Date 2022-05-13
First Publication Date 2024-07-11
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Edynak, Jérémy Philippe, Pierre
  • Souryavongsa, Eddy Keomorakott

Abstract

The present invention relates to a platform for a fan assembly, comprising a flow passage wall extending in an axial direction having an upstream end and a downstream end, and comprising two axial stiffeners. The platform comprises radial retention means positioned at the upstream end and at the downstream end of the flow passage wall, the radial retention means being designed to each define a sliding connection with a corresponding shroud of the fan assembly and in that it comprises tangential retention means designed to define a sliding pivot connection between the platform and a disc of the fan assembly.

IPC Classes  ?

  • F01D 11/00 - Preventing or minimising internal leakage of working fluid, e.g. between stages

10.

TURBOMACHINE COMPRISING A REAR INTEGRATED ELECTRICAL MACHINE

      
Application Number 18561811
Status Pending
Filing Date 2022-05-23
First Publication Date 2024-07-11
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Charier, Gilles Alain Marie
  • Truco, Romain
  • Becoulet, Julien Fabien Patrick

Abstract

A turbomachine including a low-pressure shaft, a stator supporting the low-pressure shaft by bearings including a rear bearing; an electrical machine located at a rear portion of the turbomachine and including a rotor disconnectably driven by the low-pressure shaft and a stator surrounding the rotor and attached to the stator of the turbomachine; the turbomachine includes a journal for supporting the rotor of the electrical machine, the journal including a shaft segment releasably attached to a rear end of the low-pressure shaft by extending the low-pressure shaft, a frame to which the rotor of the electrical machine is attached and which surrounds the shaft segment, and at least one bearing mounted between the frame and the shaft segment; and a sleeve that is axially movable about the low-pressure shaft and/or the shaft segment that is rotationally connected to the low-pressure shaft and the shaft segment

IPC Classes  ?

  • F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
  • F01D 25/18 - Lubricating arrangements
  • F02C 7/06 - Arrangement of bearings; Lubricating

11.

TURBOMACHINE TURBINE HAVING A CMC NOZZLE WITH LOAD ABSORPTION AND POSITIONAL ADJUSTMENT

      
Application Number 18555056
Status Pending
Filing Date 2022-04-12
First Publication Date 2024-07-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Gaillard, Aurélien
  • Cazin, Clément Emile André
  • Congratel, Sébastien Serge Francis
  • Jarrossay, Clément
  • Tabarin, Pascal Cédric

Abstract

A turbine comprising a casing and a nozzle including an outer metallic shroud (9) secured to the casing, an inner metallic shroud, and a plurality of nozzle segments (20) made of CMC forming a crown extending between the outer shroud (9) and the inner shroud, each segment including a strut (6), an inner platform, an outer platform and at least one airfoil having a hollow profile traversed by the strut (6). A turbine comprising a casing and a nozzle including an outer metallic shroud (9) secured to the casing, an inner metallic shroud, and a plurality of nozzle segments (20) made of CMC forming a crown extending between the outer shroud (9) and the inner shroud, each segment including a strut (6), an inner platform, an outer platform and at least one airfoil having a hollow profile traversed by the strut (6). For each airfoil (20), the outer platform (26) comprises an axial stop (260) extending in outward radial protrusion from the outer platform (26), and the outer metallic shroud (9) comprises a complementary axial stop (96) extending in inward radial protrusion from the outer metallic shroud (9), the axial stop (260) being upstream and axially bearing against the complementary axial stop (96), and machined with an angle of machining chosen to adjust the orientation of said at least one blade of the segment with respect to the axial direction (DA).

IPC Classes  ?

  • F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
  • F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
  • F01D 9/06 - Fluid supply conduits to nozzles or the like
  • F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings

12.

FLAME-HOLDER DEVICE FOR A TURBOJET AFTERBURNER, COMPRISING THREE-BRANCHED ARMS

      
Application Number 18555913
Status Pending
Filing Date 2022-04-15
First Publication Date 2024-07-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Bechet, Clément Jean Henri
  • Greslin, Emmanuel Jean André
  • Malbois, Pierre André Gabriel
  • Pieussergues, Christophe

Abstract

A flame-holder device for a turbojet afterburner comprises an annular row of flame-holder arms, each having an inner branch having a free end and another end, and two outer branches which extend from the other end and diverge from one another in a direction extending from the free end to the other end, such that the inner branch transitions radially outward into the two outer branches which diverge radially outward from one another in two opposing circumferential directions, thereby forming mutually approaching areas between consecutive flame-holder arms for enabling the flame to spread from arm to arm.

IPC Classes  ?

  • F23R 3/18 - Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants

13.

SYSTEM FOR CHANGING THE PITCH OF THE BLADES OF A TURBOMACHINE PROPELLER

      
Application Number 18570342
Status Pending
Filing Date 2022-06-13
First Publication Date 2024-07-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Perdrigeon, Christophe Marcel Lucien
  • Cottet, Clement

Abstract

A system for changing the pitch of the blades of a turbomachine propeller, each blade being mounted so as to pivot about a pitch axis, the system including a control means and a connection mechanism connecting the control means to each blade of the propeller, the connection mechanism including, for each blade of the propeller: a connecting rod having a first end connected to the control means; and an eccentric having a first end connected to the blade in order to rotate same about its pitch axis and a second end forming a yoke connected to a second end of the connecting rod by a pivot connection; the yoke of the eccentric being oriented towards the pitch axis of the blade.

IPC Classes  ?

  • F01D 7/00 - Rotors with blades adjustable in operation; Control thereof
  • B64C 11/32 - Blade pitch-changing mechanisms mechanical
  • F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector

14.

METHOD FOR ADDITIVE MANUFACTURING OF TURBOMACHINERY PARTS

      
Application Number 18576901
Status Pending
Filing Date 2022-06-14
First Publication Date 2024-07-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor Sistach, Hugo Jean-Louis

Abstract

A turbomachine part has a primary axis and at least one inclined portion extending in a secondary direction forming a non-zero angle with the primary axis. A method for additive manufacturing of the turbomachine part includes the steps of a) for each inclined portion: a1) providing a target roughness of an outer surface of said inclined portion, a2) providing a mechanical weakening of the inclined portion, and a3) determining a maximum roughness of the outer surface of the inclined portion according to the mechanical weakening. The method also includes steps of b) determining a total maximum roughness according to the maximum roughness of the outer surface of each inclined portion, c) determining, according to the total maximum roughness, an orientation of the primary axis of the part to be manufactured with respect to a plane of a manufacturing platen of the additive manufacturing device, and d) producing the part.

IPC Classes  ?

  • B22F 10/38 - Process control to achieve specific product aspects, e.g. surface smoothness, density, porosity or hollow structures
  • B22F 5/00 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
  • B22F 10/66 - Treatment of workpieces or articles after build-up by mechanical means
  • B22F 10/85 - Data acquisition or data processing for controlling or regulating additive manufacturing processes
  • B33Y 10/00 - Processes of additive manufacturing
  • B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
  • B33Y 50/02 - Data acquisition or data processing for additive manufacturing for controlling or regulating additive manufacturing processes
  • B33Y 80/00 - Products made by additive manufacturing

15.

TURBOFAN PROVIDED WITH A FASTENING ASSEMBLY ARRANGED ON A FAN CASING

      
Application Number 18288987
Status Pending
Filing Date 2022-04-28
First Publication Date 2024-07-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Aktir, Yasser
  • Cailliez-Tomasi, Charles

Abstract

A turbofan including a fan, at least one annular casing surrounding the fan and elements being fastened to an outer surface of the casing via a fastening assembly, the elements including a first elastic member arranged for urging the first end of the support away from the casing, and a first return member (27) arranged for resiliently urging the first end towards the casing, and a second elastic member (25) arranged for urging the second end of the support away from the casing, and a second return member (29) arranged for resiliently urging the second end towards the casing.

IPC Classes  ?

  • F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings

16.

AERONAUTICAL PROPULSION SYSTEM

      
Application Number 18541042
Status Pending
Filing Date 2023-12-15
First Publication Date 2024-06-27
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Lawniczek, Baptiste Dorian
  • Escure, Didier René André
  • Mouton, Clémentine Charlotte Marie

Abstract

A blade includes a root to be mounted in a slot of plural slots which open out at an outer periphery of a rotor disk of a movable wheel of a low-pressure turbine of an aeronautical propulsion system, the rotor disk being centered on a longitudinal axis and the slots being circumferentially distributed about the longitudinal axis, the rotor disk having an upstream surface and a downstream surface, each slot being delimited by a bottom surface, the rotor disk including blades circumferentially distributed at the outer periphery of the rotor disk, each blade being mounted in a corresponding slot, the root including a first platform which has a radially outer surface; a shroud having a second platform which has a radially inner surface; and an airfoil extending from the root to the shroud, the airfoil having a leading edge and a trailing edge.

IPC Classes  ?

  • F01D 5/30 - Fixing blades to rotors; Blade roots
  • B64C 11/04 - Blade mountings
  • B64C 11/20 - Constructional features
  • F01D 15/00 - Adaptations of machines or engines for special use; Combinations of engines with devices driven thereby

17.

AERONAUTICAL PROPULSION SYSTEM

      
Application Number 18541080
Status Pending
Filing Date 2023-12-15
First Publication Date 2024-06-27
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Lawniczek, Baptiste Dorian
  • Escure, Didier René André
  • Mouton, Clémentine Charlotte Marie

Abstract

A blade includes a root to be mounted in a slot of plural slots which opens out at an outer periphery of a rotor disk of a movable wheel of a low-pressure turbine of an aeronautical propulsion system, the rotor disk being centered on a longitudinal axis and the slots being circumferentially distributed about the longitudinal axis, the rotor disk further including blades circumferentially distributed at the outer periphery of the rotor disk, each blade being mounted in a corresponding slot, the root including a first platform which has a radially outer surface, a shroud having a second platform which has a radially inner surface, and an airfoil extending from the root to the shroud, the airfoil having a leading edge and a trailing edge. The shroud is formed of a material and has a defined difficulty.

IPC Classes  ?

18.

SYSTEME PROPULSIF AERONAUTIQUE

      
Application Number 18541057
Status Pending
Filing Date 2023-12-15
First Publication Date 2024-06-27
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Lawniczek, Baptiste Dorian
  • Escure, Didler René André
  • Mouton, Clémentine Charlotte Marie

Abstract

The invention relates to a turbine blade (33) for a movable wheel of a low-pressure turbine of an aeronautical propulsion system, comprising a root (331), a shroud (334), and an aerodynamically profiled airfoil (333) extending from the root (331) to the shroud (334), wherein a parameter N2S is defined as: The invention relates to a turbine blade (33) for a movable wheel of a low-pressure turbine of an aeronautical propulsion system, comprising a root (331), a shroud (334), and an aerodynamically profiled airfoil (333) extending from the root (331) to the shroud (334), wherein a parameter N2S is defined as: N ⁢ 2 ⁢ S = π ⁡ ( R ⁢ e 2 - Ri 2 ) × w 2 , The invention relates to a turbine blade (33) for a movable wheel of a low-pressure turbine of an aeronautical propulsion system, comprising a root (331), a shroud (334), and an aerodynamically profiled airfoil (333) extending from the root (331) to the shroud (334), wherein a parameter N2S is defined as: N ⁢ 2 ⁢ S = π ⁡ ( R ⁢ e 2 - Ri 2 ) × w 2 , where: Re is an outer flowpath radius, Ri is an inner flowpath radius, w is a limit rotation speed of the drive shaft, wherein the outer flowpath radius Re and the inner flowpath radius Ri are chosen such that N2S is greater than a certain threshold which depends on the temperature T measured when the drive shaft rotates at the limit rotation speed w.

IPC Classes  ?

  • F01D 5/30 - Fixing blades to rotors; Blade roots

19.

AERONAUTICAL PROPULSION SYSTEM

      
Application Number 18541060
Status Pending
Filing Date 2023-12-15
First Publication Date 2024-06-27
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Lawniczek, Baptiste Dorian
  • Escure, Didier René André
  • Mouton, Clémentine Charlotte Marie

Abstract

A blade includes a root to be mounted in a slot of plural slots which opens out at an outer periphery of a rotor disk of a movable wheel of a low-pressure turbine of an aeronautical propulsion system, the rotor disk being centered on a longitudinal axis and the slots being circumferentially distributed about the longitudinal axis, the root including a first platform which has a radially outer surface; a shroud including a second platform which has a radially inner surface; and an airfoil extending from the root to the shroud, the airfoil having a leading edge and a trailing edge.

IPC Classes  ?

20.

SYSTÈME PROPULSIF AÉRONAUTIQUE

      
Application Number 18541234
Status Pending
Filing Date 2023-12-15
First Publication Date 2024-06-27
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Lawniczek, Baptiste Dorian
  • Escure, Didier René André
  • Nguyen, Alexandre Thanh Nhan

Abstract

The present disclosure relates to an aeronautical propulsion system (1) comprising: a drive shaft (11); a drive turbine (8); and a fan section (2). The present disclosure relates to an aeronautical propulsion system (1) comprising: a drive shaft (11); a drive turbine (8); and a fan section (2). Figure for the abstract: FIG. 2

IPC Classes  ?

  • F01D 5/30 - Fixing blades to rotors; Blade roots
  • B64C 11/00 - Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
  • B64C 11/04 - Blade mountings
  • B64C 11/20 - Constructional features
  • F01D 25/00 - Component parts, details, or accessories, not provided for in, or of interest apart from, other groups

21.

AERONAUTICAL PROPULSION SYSTEM

      
Application Number 18541252
Status Pending
Filing Date 2023-12-15
First Publication Date 2024-06-20
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Lawniczek, Baptiste Dorian
  • Escure, Didier René André
  • Nguyen, Alexandre Thanh Nhan

Abstract

A propulsion system includes a drive shaft, a drive turbine including a drive turbine rotor having a turbine disk and turbine blades, and a fan section. The turbine disk has an inner surface, an outer surface, a bore radius, and a rim radius. One or more of the turbine blades includes a root fixed to the turbine disk in a respective slot, a stilt connected to the root and including a platform, and an airfoil connected to the platform and extending within a flowpath. The flowpath includes an outer radius and an inner radius. The drive shaft has a limit speed. The bore radius, the rim radius, the outer radius, the inner radius, and the limit speed satisfy a given relationship.

IPC Classes  ?

  • F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan

22.

FOIL FOR A TURBOMACHINE ROTOR BLADE, ASSEMBLY FOR A TURBOMACHINE ROTOR, AND TURBOMACHINE

      
Application Number 18552017
Status Pending
Filing Date 2022-03-22
First Publication Date 2024-06-20
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Goyon, Vincent
  • Mulheim, Pierre Marie
  • Martine, Arnaud
  • Giguet, Antoine
  • Delaporte, Nicolas Daniel

Abstract

A foil for a rotating blade of a turbomachine, configured to be mounted to a blade root of the rotating blade and including an overall U-shaped cross-section formed by two lateral legs connected through a lower surface, the lower surface covering a lower face of the blade root, the lateral legs each covering at least part of two lateral flanks of the blade root; and a blocking system for prohibiting at least some of relative movements between the blade root and the foil, the blocking system including at least one oblique tab, partially cut out of a lower surface of the foil and forming, with respect to the lower surface of the foil, a tilted plane protrusion able to be inserted into a housing formed in a lower face of the blade root.

IPC Classes  ?

  • F01D 5/30 - Fixing blades to rotors; Blade roots

23.

FIBROUS PREFORM WITH STIFFENERS FORMED BY UNIDIRECTIONAL YARN LAYERS

      
Application Number 18571015
Status Pending
Filing Date 2022-06-07
First Publication Date 2024-06-20
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Iglesias Cano, Celia
  • Lauwick, Lucas Antoine Christophe

Abstract

A fibrous preform of a turbomachine vane or blade includes two fibrous portions formed by three-dimensional weaving and forming an aerodynamic profile of the vane or blade and defining therebetween an interior volume, wherein each fibrous portion has at least one fibrous vibration damping stiffener including one or several unidirectional layers of yarns and present in the interior volume, and wherein the at least one fibrous stiffener is present on a non-interlinked area on which the unidirectional layer(s) of yarns are separated from the corresponding fibrous portion, and wherein the yarns of this or these unidirectional layers are woven with the corresponding fibrous portion outside the non-interlinked area, or wherein the at least one fibrous stiffener is added onto each fibrous portion.

IPC Classes  ?

  • F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
  • B29C 33/52 - SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING - Details thereof or accessories therefor with means for, or specially constructed to facilitate, the removal of articles, e.g. of undercut articles soluble or fusible
  • B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
  • B29D 99/00 - Subject matter not provided for in other groups of this subclass
  • B29K 105/08 - Condition, form or state of moulded material containing reinforcements, fillers or inserts of continuous length, e.g. cords, rovings, mats, fabrics, strands or yarns
  • B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers

24.

AERONAUTICAL PROPULSION SYSTEM

      
Application Number 18541066
Status Pending
Filing Date 2023-12-15
First Publication Date 2024-06-20
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Lawniczek, Baptiste Dorian
  • Escure, Didier René André
  • Nguyen, Alexandre Thanh Nhan

Abstract

A propulsion system includes a drive shaft, a drive turbine including a drive turbine stator and a drive turbine rotor, the rotor being connected to the drive shaft and including a turbine disk and turbine blades, and a fan section including a fan rotor and a fan stator, the drive shaft being connected to the fan rotor. The turbine disk has an inner surface, an outer surface delimiting slots distributed about a longitudinal axis, one or more of the turbine blades including a root fixed to the turbine disk in a respective slot. A bore width of the inner surface, and a rim radius of the outer surface of the turbine disk, and a limit speed of the drive shaft are defined in relation to each other.

IPC Classes  ?

  • F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
  • F01D 5/06 - Blade-carrying members, e.g. rotors - Details thereof, e.g. shafts, shaft connections
  • F01D 5/14 - Form or construction

25.

ASSEMBLY FOR A TURBINE ENGINE FAN CASE AND AN AIR INLET, METHOD FOR USING SUCH AN ASSEMBLY

      
Application Number 18555279
Status Pending
Filing Date 2022-06-02
First Publication Date 2024-06-20
Owner Safran Aircraft Engines (France)
Inventor
  • Laroche, Clement Raphael
  • Dubois, Ulrich Vincent Max
  • Minjot, Kevin Emilien Andre
  • Fameau, Vincent

Abstract

An assembly for a turbine engine fan case and an air inlet, in particular for an aircraft. The turbine engine extending along a longitudinal axis and comprising at least one fan mounted so as to be rotatable about the longitudinal axis. The fan case comprising at least a first peripheral lining. The first peripheral lining being made of an abradable material configured to be capable of being worn out during the rotation of the fan. The air inlet comprising at least one second peripheral lining. The second peripheral lining being made of an abradable material configured to be capable of being worn out during the rotation of the fan.

IPC Classes  ?

  • F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
  • F01D 7/00 - Rotors with blades adjustable in operation; Control thereof

26.

DIFFUSION CONE FOR THE REAR PART OF A JET ENGINE INCORPORATING A FLAME-HOLDER RING AT THE TRAILING EDGE

      
Application Number 18555890
Status Pending
Filing Date 2022-04-15
First Publication Date 2024-06-20
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Malbois, Pierre André Gabriel
  • Bechet, Clément Jean Henri
  • Greslin, Emmanuel Jean André
  • Lunel, Romain Nicolas

Abstract

A diffusion cone for the rear part of a jet engine has a base and, on an opposite side, a truncated tip defining an annular trailing edge, and comprises a flame-holder ring with a cross section comprising an external branch and an internal branch which are connected to one another on the base side and such that the external branch extends around the internal branch, whereby the external and internal branches delimit between them an internal space opening on the opposite side to the base. One of the branches is made up of the annular trailing edge. The flame-holder ring comprises at least one fuel inlet designed to let fuel into the internal space. Such a flame-holder ring makes it possible to ensure and to control the propagation of the flame within the afterburner channel.

IPC Classes  ?

  • F23R 3/20 - Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants incorporating fuel injection means
  • F02K 1/04 - Mounting of an exhaust cone in the jet pipe

27.

FUEL INJECTION DEVICE FOR A TURBOJET ENGINE AFTERBURNER

      
Application Number 18555930
Status Pending
Filing Date 2022-04-15
First Publication Date 2024-06-20
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor Malbois, Pierre André Gabriel

Abstract

A flame-holder device for a turbojet afterburner comprises an annular row of flame-holder arms, each having an inner branch having a free end and another end, and two outer branches which extend from the other end (52) and diverge from one another in a direction extending from the free end to the other end, such that the inner branch transitions radially outward into the two outer branches which diverge radially outward from one another in two opposing circumferential directions, thereby forming mutually approaching areas between consecutive flame-holder arms for enabling the flame to spread from arm to arm.

IPC Classes  ?

  • F23R 3/20 - Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants incorporating fuel injection means
  • F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
  • F02C 7/22 - Fuel supply systems
  • F02K 3/10 - Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluid; Control thereof by after-burners

28.

METHOD FOR MANUFACTURING A METAL ALLOY PART FOR A TURBINE ENGINE

      
Application Number 18551822
Status Pending
Filing Date 2022-03-21
First Publication Date 2024-06-13
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Turel, Arnaud
  • Voiron, Mickaël

Abstract

A method for manufacturing a metal alloy part for an aircraft turbine engine, the method including the steps of: (a) producing a blank of the part by additive manufacturing by laser fusion on a powder bed, and (b) abrasively machining the blank to obtain the part, the machining being carried out by vibratory finishing by immersing the blank in an abrasive composition contained in a tank subjected to a vibratory movement, the abrasive composition having an abrasive element formed by alumina particles, a carrying element formed by copper particles, and a liquid.

IPC Classes  ?

  • B22F 10/66 - Treatment of workpieces or articles after build-up by mechanical means
  • B22F 5/00 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
  • B22F 10/28 - Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM]
  • B22F 10/68 - Cleaning or washing
  • B23P 15/02 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from one piece
  • B33Y 10/00 - Processes of additive manufacturing
  • B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
  • B33Y 80/00 - Products made by additive manufacturing

29.

Exterior ferrule of an intermediate housing made of composite material for an aircraft turbine engine

      
Application Number 18554307
Status Pending
Filing Date 2022-04-05
First Publication Date 2024-06-13
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Vogel, Thibaut
  • Baroumes, Laurent Jean Baptiste
  • Charleux, Francois
  • Grelin, Hervé
  • Mohamed, Azath

Abstract

An exterior ferrule of an intermediate housing for an aircraft turbine engine, the ferrule made of composite material comprising a downstream annular linking end forming a peripheral groove and intended to receive a complementary linking member provided on one or more nacelle cowls, the ferrule being made from: a first fibrous preform obtained by three-dimensional weaving of yarns and densified by a matrix, the preform defining a drum of the ferrule as well as a structural part of the downstream groove wall; and a second fibrous preform obtained by three-dimensional weaving of yarns and densified by a matrix, the preform defining a base abutting the exterior of the drum defined by the first preform, as well as a structural part of the upstream groove wall.

IPC Classes  ?

  • F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
  • B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
  • B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
  • B29K 63/00 - Use of epoxy resins as moulding material
  • B29K 105/08 - Condition, form or state of moulded material containing reinforcements, fillers or inserts of continuous length, e.g. cords, rovings, mats, fabrics, strands or yarns
  • B29K 307/04 - Carbon
  • B29L 31/00 - Other particular articles

30.

METHOD FOR MANUFACTURING A HOLLOW TURBOMACHINE BLADE

      
Application Number 18554151
Status Pending
Filing Date 2022-04-04
First Publication Date 2024-06-13
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Blaise, Maxime Marie Désiré
  • Iglesias Cano, Celia
  • Lauwick, Lucas Antoine Christophe

Abstract

A method for manufacturing hollow composite material turbomachine part, includes producing a fibrous blank in one piece by three-dimensional weaving of yarns, the blank including at least one de-bonded zone forming an internal cavity, inserting a shaping part into the internal cavity of the fibrous blank in order to obtain a preform, densification of the preform by a matrix in order to obtain a structure having a fibrous reinforcement consisting of the preform densified by the matrix, wherein the shaping part is flexible and wherein the shaping part is mechanically withdrawn from the structure after the densification.

IPC Classes  ?

  • B29C 33/50 - SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING - Details thereof or accessories therefor with means for, or specially constructed to facilitate, the removal of articles, e.g. of undercut articles with means for collapsing or disassembling elastic
  • B29C 33/38 - SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING - Details thereof or accessories therefor characterised by the material or the manufacturing process
  • B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
  • B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
  • B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers

31.

FAN MODULE EQUIPPED WITH AN OIL-TRANSFER DEVICE

      
Application Number 18550228
Status Pending
Filing Date 2022-03-08
First Publication Date 2024-06-06
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Franz, Caroline Marie
  • Charier, Gilles Alain Marie
  • Maine, Herve Christian Denis
  • Morreale, Serge Rene

Abstract

Fan module for an aircraft turbine engine, the module including a fan having variable-pitch blades and an oil-transfer device configured to transfer oil between a stator and a rotor. The device includes a stator ring having internal oil ducts, a shaft engaged in the ring and having internal oil ducts, and a plain bearing and rolling bearings which are mounted between the ring and the shaft, the ring including first fluidic-connection ports which are configured to cooperate by male-female engagement with ends of oil-supply sockets during blind assembly of the device by axial translation downstream of the device as a whole.

IPC Classes  ?

  • F01D 17/16 - Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
  • F02C 7/06 - Arrangement of bearings; Lubricating

32.

EJECTION TUBE INCORPORATED IN A CENTRAL VENTILATION TUBE

      
Application Number 18551906
Status Pending
Filing Date 2022-03-22
First Publication Date 2024-05-30
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Kapala, Patrick Edmond
  • Chantoiseau, Olivier Pascal Paul

Abstract

An upstream part of an air-discharge system for a turbomachine includes: a tube with an upstream end and a downstream end, and an internal tubular structure located inside the tube and configured to form therein a narrowing of its cross section. The internal tubular structure is made in one piece with the tube at the downstream end of the tube and is freely suspended at the upstream end of the tube. A turbomachine with a device for discharging air from the bearing chambers, in which the air is conducted towards the outside of the turbomachine by the air-discharge system is provided. A method for manufacturing the air-discharge system is provided.

IPC Classes  ?

  • F01D 25/30 - Exhaust heads, chambers, or the like
  • B33Y 10/00 - Processes of additive manufacturing
  • F02C 7/06 - Arrangement of bearings; Lubricating

33.

VANE COMPRISING A STRUCTURE MADE OF COMPOSITE MATERIAL, AND ASSOCIATED MANUFACTURING METHOD

      
Application Number 18285122
Status Pending
Filing Date 2022-03-24
First Publication Date 2024-05-30
Owner Safran Aircraft Engines (France)
Inventor
  • Joudon, Vincent
  • Cottet, Clément
  • Courtier, Vivien Mickaël
  • Servant, Régis Eugène Henri

Abstract

The present invention relates to a vane (7) of a turbine engine, comprising: —a structure which has an aerodynamic profile (20); —a spar (21) comprising a vane root portion (24), a blade portion (25) arranged inside the structure which has an aerodynamic profile (20), and a stilt portion (26), the blade portion (25) comprising a body (27) connected to the vane root portion (24) and two arms extending radially from the body (27); and —a structural reinforcement (31) extending from the vane root portion (24) to the body (27) of the blade portion (25), the structural reinforcement (31) being rigidly connected to the vane root portion (24) and configured to form a force path separate from the stilt portion (26) in the event of failure of the spar (21) within the stilt portion (26).

IPC Classes  ?

  • F01D 5/14 - Form or construction
  • F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion

34.

METHOD FOR SEPARATING A PART BLANK FROM A SUPPORT BY FLUID EXPANSION IN THE ADDITIVELY MANUFACTURED SUPPORT

      
Application Number 17755466
Status Pending
Filing Date 2020-10-20
First Publication Date 2024-05-23
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Combes, Guilhem Kevin
  • Dreano, Sébastien Vincent François
  • Metge, Pierre Jean-Baptiste

Abstract

A method for separating a part blank from a support, wherein the support and the part blank are manufactured by additive manufacturing. The support includes an inner cavity containing a fluid. The separation method includes a heating of the support to deform the support by fluid expansion in the inner cavity until the support is separated from the part blank.

IPC Classes  ?

  • B22F 10/40 - Structures for supporting workpieces or articles during manufacture and removed afterwards
  • B22F 3/24 - After-treatment of workpieces or articles
  • B22F 5/00 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
  • B22F 10/64 - Treatment of workpieces or articles after build-up by thermal means
  • B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing

35.

FAN MODULE EQUIPPED WITH AN OIL TRANSFER DEVICE

      
Application Number 18549681
Status Pending
Filing Date 2022-03-09
First Publication Date 2024-05-23
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Frantz, Caroline Marie
  • Charier, Gilles Alain Marie
  • Mariot, Antoine Pierre Albert
  • Morreale, Serge René
  • Oriol, Sébastien

Abstract

A fan module for an aircraft turbine engine includes a fan with a fan shaft and variable-pitch fan blades. The fan module further includes an oil transfer device transfers oil between a stator and an actuator that controls the pitch of the fan blades the device includes a stator ring having an inner cylindrical surface and internal oil ducts, each of which opens onto the inner cylindrical surface. A shaft is engaged in the stator ring and capable of rotating rotates about an axis inside the ring. The shaft includes an outer cylindrical surface extending inside said inner cylindrical surface. The shaft has internal oil ducts, each of which opens onto said outer cylindrical surface. A rolling bearing is mounted between the ring and the shaft, on each side of the plain bearing.

IPC Classes  ?

  • F02C 7/06 - Arrangement of bearings; Lubricating
  • F16H 57/04 - Features relating to lubrication or cooling

36.

INLET CONE FOR AN AIRCRAFT TURBOMACHINE

      
Application Number 18551624
Status Pending
Filing Date 2022-03-23
First Publication Date 2024-05-23
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Louis, Alban Francois
  • Picart, Jean-Yves
  • Balland, Morgan
  • Godin, Florent Robert Andre
  • Achbari, Baghdad

Abstract

An inlet cone for an aircraft turbomachine, having at least one first portion made of a first material, referred to as rigid material, and at least one second portion made of a second material which has a hardness less than that of said first material and which is referred to as flexible material, these second portions being configured to deform elastically in the radial direction with respect to the axis when the cone is being driven to rotate.

IPC Classes  ?

  • F01D 25/02 - De-icing means for engines having icing phenomena
  • B64D 33/02 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes

37.

FAN MODULE PROVIDED WITH AN OIL TRANSFER DEVICE

      
Application Number 18550155
Status Pending
Filing Date 2022-03-03
First Publication Date 2024-05-16
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Becoulet, Julien Fabien Patrick
  • Belmonte, Olivier
  • Millier, Vincent François Georges
  • Roda, Jean Charles Olivier

Abstract

A fan module for an aircraft turbine engine includes a fan having variable pitch blades and an oil transfer device configured to provide a transfer of oil between a stator and a rotor. The oil transfer device includes a stator ring having internal oil ducts, a shaft inserted into the ring and having internal oil ducts, and an annular support of the ring configured to deform elastically to allow movements of the ring in the radial direction. A plain bearing is located between the ring and the shaft, and roller bearings are mounted between the ring and the shaft on either side of the plain bearing.

IPC Classes  ?

  • F02C 7/06 - Arrangement of bearings; Lubricating
  • F01D 7/00 - Rotors with blades adjustable in operation; Control thereof
  • F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
  • F16H 57/04 - Features relating to lubrication or cooling
  • F16H 57/08 - General details of gearing of gearings with members having orbital motion

38.

LOST WAX MOULDING MANUFACTURING PROCESS

      
Application Number 18550230
Status Pending
Filing Date 2022-03-08
First Publication Date 2024-05-16
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Guerche, Didier Maurice Marceau
  • Tami Lizuzu, Joseph Toussaint

Abstract

A method for making a wax model for the manufacture of a part, the part including at least one cavity, the method including: arranging at least two core elements made of a ceramic material into a wax mold, the core elements having at least partially a shape complementary to the cavity of the part to be manufactured, and injecting wax into the wax mold, around the core elements, so as to form the wax model, the two core elements being assembled prior to the injection of wax by at least one staple, the staple including a central portion from which two branches extend, each branch being fitted into an orifice formed in one of the core elements.

IPC Classes  ?

  • B22C 7/02 - Lost patterns
  • B22C 9/04 - Use of lost patterns
  • B22C 9/10 - Cores; Manufacture or installation of cores
  • B22C 9/22 - Moulds for peculiarly-shaped castings
  • B22C 21/14 - Accessories for reinforcing or securing moulding materials or cores, e.g. gaggers, chaplets, pins, bars
  • B22D 25/02 - Special casting characterised by the nature of the product of works of art

39.

MONITORING AN ANTI-LEAK VALVE IN A JET ENGINE

      
Application Number 18551123
Status Pending
Filing Date 2022-03-18
First Publication Date 2024-05-16
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Deneuve, Sébastien Jean Fernand
  • Verrier, Alrick Patrick Michel Jacques
  • Ricadat-Crosnier, Matthieu Louis Joseph
  • Morisot, Clémentine

Abstract

Method for monitoring an anti-leak valve of a circuit for supplying oil between an oil reservoir and elements to be lubricated of a jet engine, comprising steps of: measuring oil level values in the reservoir, at the beginning and at the end of starting the jet engine; calculating the difference between the measured oil levels, to form a measured change in the oil level; measuring an oil temperature value; determining an estimated change of the oil level in the reservoir, depending on the oil temperature; calculating the difference between the measured change and the estimated change in the oil level for the measured temperature; comparing the calculated difference with a detection threshold; and activating an alarm depending on the result of the comparison.

IPC Classes  ?

  • F16K 37/00 - Special means in or on valves or other cut-off apparatus for indicating or recording operation thereof, or for enabling an alarm to be given
  • B64D 27/16 - Aircraft characterised by the type or position of power plant of jet type
  • B64F 5/60 - Testing or inspecting aircraft components or systems

40.

DEVICE FOR CENTERING AND GUIDING A SHAFT OF AN AIRCRAFT TURBINE ENGINE

      
Application Number 18549796
Status Pending
Filing Date 2022-03-08
First Publication Date 2024-05-09
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Levisse, Paul Ghislain Albert
  • Baret, Emmanuel Fabrice Marie
  • Becoulet, Julien Fabien Patrick
  • Tan-Kim, Alexandre Jean-Marie

Abstract

A device for centering and guiding a shaft of an aircraft turbine engine, the device including an outer ring of a rolling bearing, the ring extending about an axis; an annular bearing support extending about the axis and at least partially about the ring; at least one series of connecting elements for linking the ring to the support, the connecting elements being interposed between an inner cylindrical rim of the ring and an outer cylindrical rim of the support that extends about the inner rim, and each having a first radially outer end for linking to the outer cylindrical rim and a second radially inner end for linking to the inner cylindrical rim, the ring, the support and the connecting elements being integrally formed, the connecting elements being enclosed in an annular housing that is radially delimited by the rims and laterally enclosed by an annular web.

IPC Classes  ?

  • F02C 7/06 - Arrangement of bearings; Lubricating
  • F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings

41.

Device for centering and guiding a shaft of an aircraft turbine engine

      
Application Number 18549794
Grant Number 12031448
Status In Force
Filing Date 2022-03-08
First Publication Date 2024-05-09
Grant Date 2024-07-09
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Levisse, Paul Ghislain Albert
  • Baret, Emmanuel Fabrice Marie
  • Becoulet, Julien Fabien Patrick
  • Tan-Kim, Alexandre Jean-Marie

Abstract

A device for centering and guiding a shaft of an aircraft turbine engines. The device includes an outer ring of a rolling bearing, the ring extending about a main axis and having orifices arranged around the axis axis; an annular bearing support extending about the main axis and at least partially about the ring, the support having orifices and openings arranged about the axis; and a series of studs for linking the ring to the support, the studs being distributed about the main axis and having elongation axes substantially parallel to the main axis, the body of each of the studs passing through one of the openings. Some of the studs, have bodies that pass with a first positive clearance and a second virtually zero clearance through the first openings, the first and second clearances being configured so that the device has different displacement amplitudes.

IPC Classes  ?

  • F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings

42.

TURBINE

      
Application Number 18278803
Status Pending
Filing Date 2022-02-23
First Publication Date 2024-05-09
Owner
  • SAFRAN AIRCRAFT ENGINES (France)
  • SAFRAN HELICOPTER ENGINES (France)
Inventor
  • Pellaton, Bertrand Guillaume, Robin
  • Zaia, Thibaud Louis

Abstract

A turbine for a turbomachine of longitudinal axis including an alternating arrangement of annular rows of movable blades and of fixed blades and a radially inner annular cavity formed radially inside the movable and fixed blades, and a supply circuit for supplying cooling air to the inner annular cavity, the downstream end of the supply circuit comprising an inner annular row of orifices and an outer annular row of orifices opening into the radially inner annular cavity. The turbine may also include means for controlling the flow rate of supply air to the orifices of the inner and outer annular rows of orifices.

IPC Classes  ?

  • F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
  • F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings

43.

BEARING FOR TURBOMACHINE VARIABLE PITCH STATOR VANE PIVOT, STATOR VANE COMPRISING SUCH A BEARING AND TURBOMACHINE COMPRISING SUCH STATOR VANES

      
Application Number 18548558
Status Pending
Filing Date 2022-02-23
First Publication Date 2024-05-09
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Guiot, Nicolas Thibaut
  • Frossard, Benoit André Pierre

Abstract

A bearing for a turbomachine variable pitch stator vane pivot mounted in a bore of a casing of the turbomachine and including a bushing integral with the bore and allowing rotation of a pivot rod within the casing, and additionally a ring mounted so as to be integral with the pivot rod inside the bushing and including an outer part providing the stiffening of the ring and an inner part integral with the outer part and providing a damping function.

IPC Classes  ?

  • F01D 17/16 - Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
  • F01D 25/06 - Antivibration arrangements for preventing blade vibration
  • F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings

44.

FAN MODULE EQUIPPED WITH AN OIL TRANSFER DEVICE

      
Application Number 18549415
Status Pending
Filing Date 2022-03-15
First Publication Date 2024-05-09
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Levisse, Paul Ghislain Albert
  • Frantz, Caroline Marie
  • Desombre, Didier Gabriel Bertrand
  • Perdrigeon, Christophe Marcel Lucien
  • Taillant, Jean-Claude Christian

Abstract

A fan module for an aircraft turbine engine, the module including a fan having variable-pitch blades and an oil transfer device for transferring oil between a stator and a rotor, the device including a stator ring having a cylindrical inner surface and internal oil ducts and a shaft engaged in the ring and having a cylindrical outer surface and internal oil ducts; the device includes a plain bearing located between said cylindrical inner surface and the cylindrical outer surface and defined by a single band; and rolling bearings mounted between the ring and the shaft, on either side of the plain bearing, each of these rolling bearings having one of the rings thereof integrated in one of the elements selected from the ring and the shaft.

IPC Classes  ?

  • F01D 7/02 - Rotors with blades adjustable in operation; Control thereof having adjustment responsive to speed
  • F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings

45.

DEVICE FOR CENTERING AND GUIDING A SHAFT OF AN AIRCRAFT TURBINE ENGINE

      
Application Number 18550408
Status Pending
Filing Date 2022-03-09
First Publication Date 2024-05-09
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Levisse, Paul Ghislain Albert
  • Baret, Emmanuel Fabrice Marie
  • Becoulet, Julien Fabien Patrick
  • Tan-Kim, Alexandre Jean-Marie

Abstract

A device for centering and guiding a shaft of an aircraft turbine engine is provided. The device includes an outer ring of a rolling bearing extending about an axis and having orifices, an annular bearing support extending about the axis and at least partially about the ring, the support having orifices and a series of studs for connecting the ring to the support. The studs can be distributed about the axis and extend parallel to the axis. The ends of first studs can be engaged without clearance in the orifices, and ends of second studs can be engaged with clearances in the orifices of the ring and/or of the support, the clearances being configured so that the device has different stiffnesses in at least two directions perpendicular to the axis.

IPC Classes  ?

  • F04D 29/059 - Roller bearings
  • F04D 29/66 - Combating cavitation, whirls, noise, vibration, or the like; Balancing

46.

COMPOSITE TURBOMACHINE PART FORMED OF A CORE SURROUNDED BY TWO 3D WOVEN FIBROUS PREFORMS

      
Application Number 18550879
Status Pending
Filing Date 2022-03-07
First Publication Date 2024-05-09
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Schneider-Die-Gross, Julien Paul
  • Gallet, Thomas
  • Guivarc'H, Jérémy

Abstract

A turbomachine part includes a central core which includes a first face and a second face opposite the first face, wherein the central core is surrounded firstly by a first external skin made of a composite material with a first fibrous reinforcement woven according to a three-dimensional weaving located on the first face of the central core, and secondly by a second external skin made of a composite material with a second fibrous reinforcement woven according to a three-dimensional weaving located on the second face of the central core.

IPC Classes  ?

  • F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
  • F01D 5/14 - Form or construction

47.

BLADE COMPRISING A STRUCTURE MADE OF COMPOSITE MATERIAL AND ASSOCIATED MANUFACTURING METHOD

      
Application Number 18278999
Status Pending
Filing Date 2022-02-24
First Publication Date 2024-05-02
Owner Safran Aircraft Engines (France)
Inventor Joudon, Vincent

Abstract

The present invention relates to a blade comprising: a structure of aerodynamic profile comprising two mutually opposite skins; and a spar comprising a fibrous reinforcement obtained by three-dimensional weaving and densified by the matrix, the spar comprising a blade root portion extending outside the structure of aerodynamic profile and an airfoil portion arranged inside the structure of aerodynamic profile between the two skins. Moreover, within the blade root portion, the fibrous reinforcement of the spar comprises a non-debound region and at least two debound regions extending radially from the non-debound region so as to form at least four separate branches.

IPC Classes  ?

  • F04D 29/38 - Blades
  • B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
  • B29C 70/44 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding
  • F04D 29/02 - Selection of particular materials
  • B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers

48.

TURBOJET ENGINE REAR PORTION HAVING AN AUGMENTED A9/A8 NOZZLE AREA RATIO

      
Application Number 18546543
Status Pending
Filing Date 2022-02-14
First Publication Date 2024-05-02
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Kohn, Thierry
  • Krajka, Dimitri Thomas

Abstract

A rear portion of a convergent-divergent nozzle turbojet engine includes a device for controlling the position of a divergent flap relative to that of a convergent flap, including a first connecting rod having a first end portion hinged to the convergent flap and an opposite second end portion, a second connecting rod having a first end portion hinged to the divergent flap and an opposite second end portion, and a third connecting rod having a first end portion hinged to a synchronization ring and an opposite second end portion. The second end portion of the first connecting rod is hinged to the second end portion of the second connecting rod and/or to the second end portion of the third connecting rod, and the second end portion of the third connecting rod is hinged to the second end portion of the second connecting rod.

IPC Classes  ?

  • F02K 1/12 - Varying effective area of jet pipe or nozzle by means of pivoted flaps

49.

SYSTEM AND METHOD FOR CONTROLLING THE MODIFICATION OF THE PITCH OF THE BLADES OF A TURBINE ENGINE

      
Application Number 18261702
Status Pending
Filing Date 2022-03-14
First Publication Date 2024-04-18
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor Auriol, Jacques

Abstract

A system for controlling the modification of the pitch of the blades of a fan of a turbine engine, in particular for an aircraft. The turbine engine comprising blades mounted radially in a drive shaft and a setting device configured to modify the pitch of the blades on the basis of an axial force applied to said setting device. The control system comprises a hydraulic actuator, a hydraulic pump, a connecting ring connected mechanically to the hydraulic pump such that the flow rate of the hydraulic pump is proportional to the relative speed between the drive shaft and the connecting ring, and a magnetic coupling device designed to control the drive speed of the connecting ring in order to control the pitch of the blades independently of the speed of the drive shaft.

IPC Classes  ?

  • B64C 11/40 - Blade pitch-changing mechanisms fluid, e.g. hydraulic automatic
  • F04D 27/00 - Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids

50.

AIR INLET OF A NACELLE OF AN AIRCRAFT PROPULSION ASSEMBLY FOR PROMOTING A THRUST PHASE AND A REVERSE THRUST PHASE, AND METHOD FOR USING SAME

      
Application Number 18547248
Status Pending
Filing Date 2022-02-24
First Publication Date 2024-04-18
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Cheraga, Yacine
  • Mincu, Daniel-Ciprian

Abstract

An air inlet of a nacelle of an aircraft propulsion assembly comprising an inner wall and an outer wall which are connected upstream by an air inlet lip. The air inlet lip comprising thick portions and thin portions distributed alternately over the circumference of the air inlet lip. Each thick portion comprising an upstream leading edge to separate an upstream air flow into an outer air flow guided by the outer wall and an inner air flow guided by the inner wall during a thrust phase. Each thin portion comprising a downstream leading edge situated longitudinally downstream of each upstream leading edge so as to detach a reverse air flow at the air inlet lip during a thrust reversal phase.

IPC Classes  ?

  • F02K 1/64 - Reversing fan flow
  • B64D 27/02 - Aircraft characterised by the type or position of power plant
  • F02C 7/04 - Air intakes for gas-turbine plants or jet-propulsion plants

51.

WALL PROVIDED WITH A COOLING HOLE HAVING A DIFFUSION PORTION WITH A TRIANGULAR SECTION

      
Application Number 18264847
Status Pending
Filing Date 2022-02-03
First Publication Date 2024-04-18
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Slusarz, Michel
  • Mugnier, Kevin Michael

Abstract

A turbomachine component such as a blade, includes a wall provided with at least one cooling hole having a metering portion and a tapered diffusion portion. The diffusion portion includes a central edge forming a bottom of the diffusion portion that is oblique with respect to a flow axis of the cooling hole.

IPC Classes  ?

  • F01D 5/18 - Hollow blades; Heating, heat-insulating, or cooling means on blades
  • F01D 25/12 - Cooling
  • F23R 3/00 - Continuous combustion chambers using liquid or gaseous fuel

52.

VENTILATION RING FOR THE REAR BEARING SUPPORT MEMBER OF AN AIRCRAFT TURBINE ENGINE

      
Application Number 18273138
Status Pending
Filing Date 2022-01-20
First Publication Date 2024-04-18
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Voiron, Mickaël
  • Garnier, Fabien Stéphane
  • Ovaere, Nicolas

Abstract

A ventilation ring for a bearing support member of an aircraft turbine engine includes two or more tubular walls and one or more spacer walls. The two or more tubular walls extend opposite each other, and one or more of the tubular walls includes a main portion and a base, the base having a thickness greater than a thickness of the main portion. The one or more spacer walls connect the two or more tubular walls. The one or more spacer walls have one or more apertures extending into the base and a rim around the aperture. An entirety of the rim has a thickness greater than a thickness of the spacer wall at a distance from the rim.

IPC Classes  ?

  • F01D 25/12 - Cooling
  • F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings

53.

PROPULSION ASSEMBLY FOR AN AIRCRAFT COMPRISING A STATOR VANE INTEGRATED INTO AN UPSTREAM PART OF A MOUNTING PYLON OF REDUCED HEIGHT

      
Application Number 18261124
Status Pending
Filing Date 2022-01-05
First Publication Date 2024-04-11
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Lebeault, Eva Julie
  • Binder, Anthony
  • Soulat, Laurent

Abstract

A propulsion assembly for an aircraft comprising a dual-flow turbine engine equipped with a fan, an aerodynamic outer shroud acting as a nacelle as well as a mounting pylon, the propulsion assembly having a secondary flow path defined by an outer radial defining surface formed by the shroud, the turbine engine including stator vanes, and the mounting pylon comprising a part housed in the secondary flow path, referred to as upstream part. According to the invention, the upstream part of the pylon extends radially from the inner radial defining surface, along a radial pylon height strictly less than a total radial height of the secondary flow path, and the upstream part of the pylon extends in the downstream direction from a root part of one of the stator vanes.

IPC Classes  ?

  • B64D 27/40 - Arrangements for mounting power plants in aircraft
  • B64D 27/12 - Aircraft characterised by the type or position of power plant of gas-turbine type within, or attached to, wing
  • F02C 7/20 - Mounting or supporting of plant; Accommodating heat expansion or creep
  • F02K 1/64 - Reversing fan flow

54.

TURBINE ROTOR FOR A TURBOMACHINE

      
Application Number 18271893
Status Pending
Filing Date 2022-01-18
First Publication Date 2024-04-11
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Girardeau, Julian Nicolas
  • Silet, Benoit Guillaume

Abstract

A rotor of a turbine extending around a longitudinal axis includes a rotor disc bearing at its periphery one or more rotor blades, each blade including an airfoil provided at its radially outer end with a platform equipped with an upstream lip and a downstream lip, the platform having a front lateral edge and a back lateral edge with respect to the normal direction of rotation of the rotor about the longitudinal axis, and a suction face portion which extends from the front lateral edge to the suction face of the airfoil. The upstream lip and the downstream lip each includes sawtooth-shaped cut-outs such as to form a front tooth and a back tooth which are contiguous, in that the back tooth has a cutting edge and in that the front tooth of each lip is disposed on the suction face portion of the platform.

IPC Classes  ?

  • F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
  • F01D 5/20 - Specially-shaped blade tips to seal space between tips and stator

55.

METHOD FOR REPAIRING AN END PLATE OF A TURBOMACHINE ROTOR

      
Application Number 18264969
Status Pending
Filing Date 2022-02-09
First Publication Date 2024-04-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Lexilus, Jean-Hilaire
  • Bot, Cédric Nicolas
  • Reghezza, Patrick Jean-Louis

Abstract

A method for repairing an end plate of a turbomachine rotor, the end plate including a crown having at least two collar fastening holes and at least one balancing-weight fastening hole located between the two collar fastening holes, each collar fastening hole being equipped with a crimped nut, the crown having a damaged portion between two collar fastening holes, the method including the steps of removing the two crimped nuts located on either side of the damage; removing the damaged crown portion, for example by machining; putting in place a strip in the form of a crown portion closing off the removed crown portion, the strip having at least one balancing-weight fastening hole and two mounting holes; fastening the strip to the crown with two crimped nuts, which penetrate the mounting holes, by crimping each nut in a collar fastening hole.

IPC Classes  ?

  • F01D 5/02 - Blade-carrying members, e.g. rotors
  • B23P 6/00 - Restoring or reconditioning objects
  • F01D 5/30 - Fixing blades to rotors; Blade roots

56.

METHOD FOR PREDICTING THE WEAR OF A MECHANICAL PART, AND METHOD FOR DETERMINING THE OPERATIONAL RISK ASSOCIATED WITH AN AIRCRAFT OR WITH A FLEET OF AIRCRAFT

      
Application Number 18553952
Status Pending
Filing Date 2022-04-05
First Publication Date 2024-04-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Coupard, Josselin Xavier
  • Bonnet, Guillaume Rémi
  • Gaudart, Germain

Abstract

A method for predicting the wear of a mechanical part and uncertainty of this prediction in a profile of use, the profile of use taking into account environmental conditions associated with environmental data and the time of use of the mechanical part under each of these environmental conditions, the method including determining operational data associated with a plurality of mechanical parts of the same type as the mechanical part; on the basis of the operational data determined, determining a plurality of predictive models of the wear of the part, the mechanical part being able to be divided into a plurality of elements, each element being modelled using at least one model of the plurality of models; and, for each model, determining a weighting coefficient, determining a wear prediction of the mechanical part, and determining a statistical quantity representative of dispersion of the predictions of the plurality of models.

IPC Classes  ?

  • F01D 21/00 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
  • B64F 5/60 - Testing or inspecting aircraft components or systems

57.

TURBINE STATOR ASSEMBLY WITH A RADIAL DEGREE OF FREEDOM BETWEEN A GUIDE VANE ASSEMBLY AND A SEALING RING

      
Application Number 18256339
Status Pending
Filing Date 2021-12-03
First Publication Date 2024-03-28
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Sultana, Patrick Jean Laurent
  • Got, Marc-Antoine Anatole
  • Zamai, Laurent Cédric

Abstract

A turbine stator assembly including a guide vane assembly and a sealing ring bearing an abradable element of a dynamic sealing ring. The assembly includes pins secured to the sealing ring and cooperating with respective oblong openings formed in the guide vane assembly so as to allow a radial movement of the ring with respect to the guide vane assembly in order to compensate for differential thermal expansions. A seal is arranged so as to compensate for circumferential spaces between ring sectors forming the guide vane assembly.

IPC Classes  ?

  • F01D 11/00 - Preventing or minimising internal leakage of working fluid, e.g. between stages
  • F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector

58.

REDUCED BULK ACOUSTIC TREATMENT PANEL FOR A TURBOJET ENGINE

      
Application Number 18257810
Status Pending
Filing Date 2021-12-14
First Publication Date 2024-03-28
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Riou, Georges Jean Xavier
  • Gea Aguilera, Fernando

Abstract

An acoustic treatment panel including acoustic absorption cells each comprising a bottom wall, a porous inlet wall, an enclosure extending in an axial direction between the bottom wall and the inlet wall, and a first acoustic horn extending inside the enclosure between a first opening and a second opening that is smaller than said first opening, the first opening facing said inlet wall. Each cell comprises a second horn extending inside the enclosure in the first axial direction between a first opening of the second horn and a second opening of the second horn that is smaller than said first opening.

IPC Classes  ?

  • F02K 1/34 - Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto using fluid jets to influence the jet flow for attenuating noise
  • F02C 7/045 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for noise suppression

59.

METHOD AND SYSTEM FOR DETERMINING AIRCRAFT LANDING RUNWAY CONDITIONS

      
Application Number 18265103
Status Pending
Filing Date 2021-12-01
First Publication Date 2024-03-28
Owner
  • SAFRAN AIRCRAFT ENGINES (France)
  • SAFRAN ELECTRONICS & DEFENSE (France)
  • SAFRAN LANDING SYSTEMS (France)
  • SAFRAN NACELLES (France)
Inventor
  • Maalioune, Hakim
  • Miralles, Laurent
  • Bastide, Christophe
  • Fert, Jérémy Edmond
  • Hupin, Vincent

Abstract

Said method for determining aircraft landing runway conditions comprises the steps of: acquiring a set of data groups of different types (D1, D2) for evaluating and monitoring runway degradation conditions; deriving weighting coefficients (Ki) from each data group; filtering the data; determining, for each data group, a partial runway condition; modifying the weighting coefficients of each data group; and combining the partial runway conditions to derive a runway condition coefficient (RWYCC) associated with a confidence index (C1) derived from the modified weighting coefficients.

IPC Classes  ?

  • G08G 5/02 - Automatic landing aids, i.e. systems in which flight data of incoming planes are processed to provide landing data
  • G08G 5/00 - Traffic control systems for aircraft

60.

SEALING DEVICE FOR A DISCHARGE VALVE OF A TURBOMACHINE

      
Application Number 18275210
Status Pending
Filing Date 2022-01-21
First Publication Date 2024-03-28
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Mohamed, Azath
  • Branco, Alexandre
  • Papin, Thierry Georges Paul
  • Collin, Virginie Martine

Abstract

Discharge valve of an aircraft turbomachine, including a discharge gate intended to be mounted pivotably about an axis between a position for closing an air passage orifice and an open position of said orifice, the orifice being formed in an intermediate casing of the turbomachine, the discharge valve including a sealing device having a seal configured to be movably mounted on the intermediate casing and a movement system configured to move the seal between a first position in which the seal is intended to bear against the discharge gate when the gate is in a closure position, and a second position in which the seal is retracted with respect to the air passage orifice when the discharge gate is in an open position.

IPC Classes  ?

  • F02C 9/18 - Control of working fluid flow by bleeding, by-passing or acting on variable working fluid interconnections between turbines or compressors or their stages
  • F02C 7/28 - Arrangement of seals
  • F16K 3/10 - Gate valves or sliding valves, i.e. cut-off apparatus with closing members having a sliding movement along the seat for opening and closing with flat sealing faces; Packings therefor with pivoted closure members with special arrangements for separating the sealing faces or for pressing them together

61.

IMPROVED FERRULE FOR COUNTER-ROTATING TURBINE IMPELLER

      
Application Number 17754858
Status Pending
Filing Date 2020-10-08
First Publication Date 2024-03-21
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Cousseau, Simon Jean-Marie Bernard
  • Delalandre, Mathieu Patrick Henri
  • Sultana, Patrick Jean Laurent
  • Zamai, Laurent Cédric

Abstract

A counter-rotating turbine for a turbine engine comprising an inner rotor having an inner drum to which is attached a plurality of inner impellers rotatably supported by a first shaft, an outer rotor comprising an outer drum to which is fastened a plurality of outer impellers rotatably supported by a second shaft coaxial with the first shaft, the outer rotor comprising a downstream impeller having a plurality of downstream moving blades extending between an outer shroud and an inner shroud, one upstream end of the outer shroud being attached downstream of said outer drum, the inner shroud being attached to the second shaft, at least one of the inner shroud and of the outer shroud comprising at least one flexible transition part configured to allow elastic deformation of said shroud in the radial direction.

IPC Classes  ?

  • F02C 3/067 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages having counter-rotating rotors

62.

METHOD FOR MANUFACTURING A VANE MADE OF A COMPOSITE MATERIAL WITH INTEGRATED ATTACHMENT LUGS AND PLATFORMS

      
Application Number 18254972
Status Pending
Filing Date 2021-11-22
First Publication Date 2024-03-21
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Masson, Antoine Hubert Marie Jean
  • Iglesias Cano, Célia
  • Laborieux, Larry Sylvio

Abstract

A method for manufacturing a turbomachine vane made of composite material, the method including forming, by three-dimensional or multilayer weaving, a fiber blank separated across its thickness into first, second and third parts in two non-interlinked areas present at the longitudinal ends of the fiber blank, the first part located between the second and third parts to which it is connected by weaving outside non-interlinked areas, forming, from the fiber blank, a preform of the vane to be produced, by unfolding and shaping, at each longitudinal end and on either side of the first part, the segments of the second part and the segments of the third part not interlinked with the first part to form preform parts for a platform of the vane, the segments of the first part not interlinked with the segments of the second and third parts extending along the longitudinal axis, and densifying the preform.

IPC Classes  ?

  • F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion

63.

OIL JET FOR A TURBINE ENGINE

      
Application Number 18256816
Status Pending
Filing Date 2021-12-06
First Publication Date 2024-03-21
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor Picard, Mathieu

Abstract

An oil jet for lubricating members such as bearings inside a turbine engine, includes a main pipe for supplying oil, at least one secondary oil discharge pipe through which the oil is sprayed onto the member, and a connection area which is for connecting the secondary pipe to the main pipe and which includes a junction angle between the main pipe and the secondary pipe, wherein the connection area includes an inner wall in contact with the oil, the inner wall being at least partially curved at the junction between the secondary pipe and the main pipe such that the junction angle inside the connection area is at least partially rounded.

IPC Classes  ?

  • F01D 25/18 - Lubricating arrangements
  • F16C 33/66 - Special parts or details in view of lubrication

64.

INNER BLADE SUPPORT RING OF A TURBOMACHINE COMPRESSOR STATOR

      
Application Number 18274645
Status Pending
Filing Date 2022-02-02
First Publication Date 2024-03-21
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor Cornec, Nicolas Gérard Bénito

Abstract

An inner ring for supporting variable-pitch blades of an axial compressor stator of a turbomachine includes two or more ring sectors, assembled to form the inner ring, each ring sector including plural housings for receiving a pivot axis of one of the blades. At each junction plane between the ends of two successive ring sectors, an assembly pin and a cavity for receiving the assembly pin are engaged. The assembly pin is fixed on one end one of the two successive ring sectors and the cavity is formed at the end of the other of the two successive ring sectors. The cavity receives the assembly pin with a mounting clearance allowing a mounting by sliding fit, and the assembly pin is made of a material having a coefficient of thermal expansion greater than a coefficient of thermal expansion of a material constituting the successive ring sectors.

IPC Classes  ?

  • F01D 17/16 - Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
  • F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
  • F04D 19/02 - Multi-stage pumps
  • F04D 29/52 - Casings; Connections for working fluid for axial pumps
  • F04D 29/54 - Fluid-guiding means, e.g. diffusers
  • F04D 29/56 - Fluid-guiding means, e.g. diffusers adjustable

65.

Hybridization of the fibers of the fibrous reinforcement of a fan blade

      
Application Number 18039611
Grant Number 12000303
Status In Force
Filing Date 2021-11-29
First Publication Date 2024-03-21
Grant Date 2024-06-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Rakotoarisoa, Carole Onja
  • Fixy, Teddy
  • Gondre, Guillaume Pascal Jean-Charles
  • Schneider-Die-Gross, Julien Paul

Abstract

The invention relates to a fan blade (3) made of composite material the fibrous reinforcement of which comprises first strands (12) having a first stiffness, second strands (13) having a second stiffness and third strands (14) having a third stiffness lower than the first stiffness and greater than the second stiffness, the blade comprising a first portion (15) comprising only first strands (12), a second portion (16) comprising only second strands (13) and a third portion (17) which is located between the first and second portions (15, 16) and which comprises both first, second and third strands (12, 13, 14); there is a gradual transition of properties at the interface between the different portions (15, 16, 17).

IPC Classes  ?

  • F01D 5/14 - Form or construction
  • F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion

66.

TURBINE RING ASSEMBLY MOUNTED ON A CROSS-MEMBER

      
Application Number 18555801
Status Pending
Filing Date 2022-04-12
First Publication Date 2024-03-21
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Lemonnier, Jérôme Claude George
  • Boisnault, Franck Davy
  • Gaudry, Florian Armand Gabriel
  • Baudy, Kévin Claude Luc
  • Chambre, Marion France

Abstract

A turbine ring assembly extending around an axis, including a plurality of ring sectors made of ceramic matrix composite material forming a turbine ring and a ring support structure held by a turbine casing, each ring sector including a base from which an upstream tab and a downstream tab extend radially outward, spaced axially from one another, wherein the assembly also includes a cross-member mechanically connected to the ring support and including a first radial flange bearing against the upstream latching tabs of the ring sectors, and, for each ring sector, at least two transverse pins and a radial spring, each transverse pin passing through the upstream latching tab and the downstream latching tab of the ring sector and the ring support to hold the ring sector and the ring support secured to one another.

IPC Classes  ?

  • F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
  • F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part

67.

ADJUSTMENT OF TWIST TO OPTIMIZE MOULDS AND THE SHAPING OF TEXTILE ARCHITECTURES

      
Application Number 18554388
Status Pending
Filing Date 2022-03-28
First Publication Date 2024-03-14
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Wielhorski, Yanneck
  • Schneider-Die-Gross, Julien Paul

Abstract

An installation for compacting strands or a preform including strands, the strands or the preform including at least a first portion in which the strands are twisted at a first angle and a second portion in which the strands are twisted at a second angle different from the first angle, the installation including a mould which includes the strands or the preform, and wherein the mould has at least a first part in which the first portion of the strands or of the preform is disposed, having a first thickness and a second part in which the second portion of the strands or of the preform is disposed, having a second thickness different from the first thickness.

IPC Classes  ?

  • D07B 7/02 - Machine details; Auxiliary devices
  • B29C 43/02 - Compression moulding, i.e. applying external pressure to flow the moulding material; Apparatus therefor of articles of definite length, i.e. discrete articles
  • B29C 70/20 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in a single direction, e.g. roving or other parallel fibres

68.

METHOD AND DEVICE FOR CHECKING THE COMPLIANCE OF A WORKPIECE

      
Application Number 17754585
Status Pending
Filing Date 2020-10-07
First Publication Date 2024-03-07
Owner
  • SAFRAN AIRCRAFT ENGINES (France)
  • SAFRAN (France)
Inventor
  • Descoqs, Augustin Marie Michel
  • Ricordeau, Julien Alexis Louis
  • Sulli, Quentin

Abstract

Method for checking the compliance of a workpiece, comprising the following steps: estimating a risk of non-compliance of characteristic on the basis of a law of probability associated with the characteristic; and verifying whether the estimated risk of non-compliance satisfies a decision criterion and, if so, declaring that the workpiece is compliant for characteristic; if not, measuring a value of the characteristic, determining whether or not the workpiece is compliant based on the measured value, and updating the law of probability associated with characteristic based on the measured value.

IPC Classes  ?

  • G05B 19/418 - Total factory control, i.e. centrally controlling a plurality of machines, e.g. direct or distributed numerical control (DNC), flexible manufacturing systems (FMS), integrated manufacturing systems (IMS), computer integrated manufacturing (CIM)

69.

TURBINE RING ASSEMBLY FOR A TURBOMACHINE

      
Application Number 18552716
Status Pending
Filing Date 2022-03-25
First Publication Date 2024-02-29
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Jarrossay, Clément
  • Gaillard, Aurélien
  • Tabarin, Pascal Cédric
  • Nimhauser, Arthur Paul Gabriel
  • Cazin, Clément Emile André

Abstract

A turbine ring assembly having ring segments made of ceramic matrix composite material and each having first and second attachment tabs and a cavity for the circulation of air flow, a metal support having a first bracket and a second bracket bearing axially upstream against the second tab, a first metal flange arranged upstream of the first bracket and having an inner periphery bearing axially downstream against the first tab and an outer periphery fastened to the first bracket, and air passage orifices formed in the inner periphery of the first flange and/or in the second bracket, the orifices configured to ensure that the air flow passes from the cavity to the outside of the assembly.

IPC Classes  ?

  • F01D 11/24 - Actively adjusting tip-clearance by selectively cooling or heating stator or rotor components
  • F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings

70.

HYBRID TURBOFAN ENGINE FOR AN AIRCRAFT, COMPRISING A MOTOR/GENERATOR COOLED BY HEAT PIPES

      
Application Number 18261443
Status Pending
Filing Date 2022-01-07
First Publication Date 2024-02-29
Owner
  • SAFRAN (France)
  • SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Sibilli, Thierry
  • Millier, Vincent François Georges
  • Guillotel, Loïc Paul Yves

Abstract

A hybrid turbofan engine for an aircraft, comprising a fan, an electric motor/generator and a gas generator, the engine comprising, between the fan and a splitter that separates the flows, an internal wall for delimiting an air duct, situated upstream of inlet guide vanes of a primary duct, and, upstream of outlet guide vanes, an internal upstream wall for delimiting a secondary duct, the motor/generator having a stator carried by a stator support fixed to a stator part of the engine. According to the invention, a plurality of heat pipe for cooling the motor/generator are provided, each heat pipe having an evaporation section fixed to the stator support, and a condensation section fixed to the internal wall or the internal upstream wall.

IPC Classes  ?

  • F01D 15/10 - Adaptations for driving, or combinations with, electric generators
  • B64D 27/24 - Aircraft characterised by the type or position of power plant using steam, electricity, or spring force
  • F02C 7/12 - Cooling of plants
  • F02K 3/115 - Heating the by-pass flow by means of indirect heat exchange

71.

Method for manufacturing a composite platform for a fan of an aircraft turbine engine

      
Application Number 17908500
Grant Number 11939937
Status In Force
Filing Date 2021-03-01
First Publication Date 2024-02-22
Grant Date 2024-03-26
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Minervino, Matteo
  • Fromonteil, Didier
  • Grelin, Hervé
  • Lamouche, Damien Bruno

Abstract

b), on which a fastening tab is located (34), said fastening tab being configured to be fixed to a fan disc (2). The invention is characterized in that it comprises the steps of: a) preparing fabrics or sheets which are pre-impregnated with a resin, b) depositing the fabrics or sheets in a mold, c) positioning a metal reinforcement (36) in the mold on the fabrics or sheets, the reinforcement being integrally formed with said fastening tab, d) depositing the fabrics or sheets on a part of the reinforcement, and e) closing and heating the mold for solidification of the assembly formed by the fabrics or sheets and the reinforcement.

IPC Classes  ?

  • F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
  • B29C 70/72 - Encapsulating inserts having non-encapsulated projections, e.g. extremities or terminal portions of electrical components
  • B29K 705/00 - Use of metals, their alloys or their compounds, for preformed parts, e.g. for inserts
  • B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers

72.

IMPROVED ACOUSTIC ATTENUATION DEVICE FOR AN AIRCRAFT PROPULSION UNIT

      
Application Number 18260967
Status Pending
Filing Date 2022-01-07
First Publication Date 2024-02-22
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Gea Aguilera, Fernando
  • Dubois, Adrien Clément Marcel
  • Boiteux, Jean-Michel Daniel Paul
  • Wiart, Ludovic Mathieu René

Abstract

Propulsion assembly for an aircraft comprising a turbomachine having at least one rotating part rotating about an axis of rotation, an attachment strut, and a structural element carrying the turbomachine via the attachment strut, the rotating part being disposed upstream of the structural element and of the attachment strut such that an air jet emerging from the rotating part, in the wake of thereof, impacts the structural element and the attachment strut, a leading edge of the structural element and/or of said attachment strut locally comprising at least one acoustic attenuation device disposed at least partly in the wake of the rotating part, the acoustic attenuation device being a local modification of the structure and/or of the profile of the leading edge.

IPC Classes  ?

  • B64C 11/00 - Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft

73.

DEVICE FOR SHAPING A FIBROUS PREFORM FOR PRODUCING A BLADED PART OF A TURBOMACHINE

      
Application Number 18557003
Status Pending
Filing Date 2022-04-22
First Publication Date 2024-02-22
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Blaise, Maxime Marie Desire
  • Fabre, Hubert Jean Marie
  • Pote, Jeff

Abstract

A device for shaping at least one fibrous preform of a bladed part of a turbine engine, the device including a mould formed of multiple parts nested inside one another, the mould defining an internal cavity for enclosing the preform entirely, the cavity having two platform zones and a blade zone extending between the two platform zones, wherein the mould includes at least a lower shell, an upper shell, a side shell, and end shells, and in that each of the side and end shells includes three elements, respectively lower, intermediate and upper.

IPC Classes  ?

  • B29D 99/00 - Subject matter not provided for in other groups of this subclass

74.

DEVICE FOR MOULDING A BLADED PART OF A TURBOMACHINE

      
Application Number 18557002
Status Pending
Filing Date 2022-04-22
First Publication Date 2024-02-22
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Blaise, Maxime Marie Desire
  • Fabre, Hubert Jean Marie
  • Pote, Jeff

Abstract

A device for moulding at least one bladed part of a turbine engine, including a base; a mould formed from a plurality of parts nested inside one another, this mould being applied to the base. A first seal is mounted between the mould and the base; and a bell is mounted on the mould and around the mould, this bell being applied to the base. A second seal is mounted between the bell and the base, this bell being configured to be held tight against the base and having interior surfaces engaging by wedge effect with complementary exterior surfaces of the mould in order to apply a clamping force on the parts of this mould.

IPC Classes  ?

  • B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
  • B29C 33/00 - SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING - Details thereof or accessories therefor
  • B29C 61/06 - Making preforms having internal stresses, e.g. plastic memory
  • B29C 70/54 - Component parts, details or accessories; Auxiliary operations
  • F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion

75.

AIRCRAFT TURBINE ENGINE ASSEMBLY COMPRISING A HOLDER FOR FITTINGS

      
Application Number 18257830
Status Pending
Filing Date 2021-12-15
First Publication Date 2024-02-15
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Beutin, Bruno Albert
  • Delaforge, Anthony Pierre Guy
  • Zsiga, Zoltan

Abstract

Assembly for an aircraft turbine engine, comprising an intermediate casing hub, an accessory gearbox, and a transfer case including a power transmission shaft configured to transmit mechanical power, taken off an engine shaft of the turbine engine, to at least one accessory mounted on the accessory gearbox, in which the transfer case comprises a structural element surrounding the power transmission shaft and rigidly attaching the accessory gearbox to the intermediate casing hub.

IPC Classes  ?

  • F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
  • F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
  • F02C 7/20 - Mounting or supporting of plant; Accommodating heat expansion or creep
  • F02C 3/10 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor

76.

TURBOMACHINE MODULE EQUIPPED WITH AN ELECTRIC MACHINE, AND TURBOMACHINE EQUIPPED WITH SUCH A MODULE

      
Application Number 18259226
Status Pending
Filing Date 2021-12-20
First Publication Date 2024-02-15
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Dautreppe, Frédéric
  • Galivel, Jean-Pierre Elie

Abstract

An aircraft turbomachine module, including an accessory gearbox, a transfer shaft, and a drive shaft rotatably connected with one another by a power transmission device. The power transmission device can be housed in a gearbox and the module can further include an electric machine rotatably connected to the transfer shaft. The electric machine can be interposed between the transmission housing and at least one part the transfer shaft.

IPC Classes  ?

  • F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
  • F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
  • F02C 7/275 - Mechanical drives
  • F02C 7/06 - Arrangement of bearings; Lubricating

77.

ASSEMBLY COMPRISING AN AIRCRAFT TURBINE ENGINE AND MOUNTING PYLON THEREOF

      
Application Number 18260923
Status Pending
Filing Date 2022-01-13
First Publication Date 2024-02-15
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Glemarec, Guillaume
  • Vignes, Jean-Baptiste
  • Gentils, Romuald Muriel
  • Capasso, Valerio

Abstract

Assembly including an aircraft turbine engine and a pylon for mounting the turbine engine to an element of the aircraft. The pylon includes members for suspending the turbine engine, the members being connected to the turbine engine in at least one plane which is perpendicular to the axis and which is located upstream of the combustion chamber of the turbine engine, such that the turbomachine is cantilevered to the pylon, The assembly further includes at least one damper which connects the turbomachine to the pylon and which is located in a plane perpendicular to the axis located downstream of the combustion chamber, the damper being configured to limit the relative movements between the turbine engine and the pylon without transmitting force.

IPC Classes  ?

  • F02C 7/20 - Mounting or supporting of plant; Accommodating heat expansion or creep
  • F01D 25/04 - Antivibration arrangements
  • B64D 27/26 - Aircraft characterised by construction of power-plant mounting

78.

METHOD FOR PROCESSING VOLUME IMAGES BY PRINCIPAL COMPONENT ANALYSIS

      
Application Number 18259488
Status Pending
Filing Date 2021-12-24
First Publication Date 2024-02-15
Owner
  • SAFRAN (France)
  • SAFRAN AIRCRAFT ENGINES (France)
  • ECOLE NORMALE SUPERIEURE PARIS- SACLAY (France)
  • CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
Inventor
  • Schneider-Die-Gross, Julien Paul
  • Jailin, Clément
  • Mendoza Quispe, Arturo
  • Roux, Stéphane

Abstract

Method for processing a plurality of X-ray tomography volume images each associated with a part, the plurality of volume images comprising a reference volume image, including: a step of correlating volume images to obtain a displacement field between each image and the reference image, to obtain a plurality of displacement fields minimizing the difference between the volume images, a processing by a dimensionality reduction method of the plurality of the image displacement fields to express them according to eigenmodes, and a statistical analysis of the fields expressed according to the eigenmodes.

IPC Classes  ?

79.

AUXILIARY OIL TANK FOR AN AIRCRAFT TURBINE ENGINE

      
Application Number 18491888
Status Pending
Filing Date 2023-10-23
First Publication Date 2024-02-08
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Jacquemard, Christophe Paul
  • Desombre, Didier Gabriel Bertrand

Abstract

An auxiliary tank for an aircraft turbine engine is provided, and in conjunction with a pump and associated auxiliary lubrication circuit, supplies oil to a reducer when, for example, a phase of free rotation of the fan is detected. Thus, the reducer is always lubricated, even during the phases of free rotation of the fan, thus ensuring a longer life of the reducer gears. The auxiliary tank is arranged with respect to the reducer so that oil is recovered from the reducer at least in part by the auxiliary tank.

IPC Classes  ?

  • F02C 7/06 - Arrangement of bearings; Lubricating
  • B64D 27/02 - Aircraft characterised by the type or position of power plant
  • F01D 25/18 - Lubricating arrangements
  • F16H 57/04 - Features relating to lubrication or cooling

80.

GUIDE VANE ASSEMBLY FOR AN AIRCRAFT TURBINE ENGINE

      
Application Number 18255351
Status Pending
Filing Date 2021-11-26
First Publication Date 2024-02-01
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor Merlin, Remi

Abstract

A guide vane assembly of an aircraft turbine engine, the guide vane assembly being segmented into a plurality of segments arranged circumferentially next to one another about an axis. The guide vane assembly includes two annular rings connected together by vanes and the geometry of the inner ring is modified to reduce the vibrational responses of the guide vane assembly

IPC Classes  ?

  • F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector

81.

TURBINE BLADE FOR AN AIRCRAFT TURBOMACHINE, PROVIDED WITH A CHANNEL FOR EJECTING A PRIMARY FLOW TOWARDS AN INTER-LIP CAVITY

      
Application Number 18256317
Status Pending
Filing Date 2021-12-01
First Publication Date 2024-02-01
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Coussirou, Jean Charles Marie
  • Langevin, Thomas
  • Silet, Benoit Guillaume

Abstract

A blade and a sealing element forming an assembly for a turbomachine turbine. The blade includes an airfoil and a platform to which the sealing element is connected. An internal channel passes through the platform so as to draw off fluid circulating in a primary duct that it delimits and to inject the thus drawn-off fluid into an inter-lip cavity delimited by the sealing element.

IPC Classes  ?

  • F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
  • F01D 5/14 - Form or construction
  • F01D 11/10 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using sealing fluid, e.g. steam

82.

TURBOMACHINE ROTOR HAVING IMPROVED VIBRATORY BEHAVIOUR

      
Application Number 18264343
Status Pending
Filing Date 2022-02-01
First Publication Date 2024-02-01
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Phouratsamay, Sylvain
  • Avakian, Sylvain Frédéric
  • Cavarec, Mickael

Abstract

Assembly comprising a body extending around a central axis, the body having an outer surface from which a plurality of blades extend, each of said blades having a blade root and a blade tip, defining an inner radial end and an outer radial end of the blade relative to the central axis, said blades having the same blade height measured radially relative to the central axis, characterized in that each of the blades is connected to the body by its blade root via a connection having a nonzero connection height, so that for the plurality of said blades, the connection height of two successive blades is different.

IPC Classes  ?

  • F01D 5/16 - Form or construction for counteracting blade vibration

83.

TURBINE STATOR ASSEMBLY

      
Application Number 18547997
Status Pending
Filing Date 2022-03-04
First Publication Date 2024-01-25
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Danteny, Ulysse Jacques Bernard
  • Beauboit, Kenny

Abstract

The invention relates to a stator assembly for an aircraft turbomachine extending about an axis and including: a shroud ring, a plurality of bladed ring sectors, sliding assembly allowing relative movement between the shroud ring and each of the ring sectors, the sliding assembly including two outer annular flanges which are upstream and downstream of the shroud ring, respectively, and an inner flange of each of the ring sectors which is inserted axially between the two outer flanges, wherein each of the ring sectors includes a deflector which is supported by the inner flange of the sector and extends around the upstream annular flange of the shroud ring.

IPC Classes  ?

  • F01D 11/00 - Preventing or minimising internal leakage of working fluid, e.g. between stages
  • F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector

84.

Method for manufacturing a turbomachine compressor blade by compacting

      
Application Number 18030425
Grant Number 11904420
Status In Force
Filing Date 2021-09-30
First Publication Date 2024-01-18
Grant Date 2024-02-20
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Salmon, Jérôme
  • Gimel, Alexandre
  • Gleize, Christophe
  • Pereira, Axel
  • Richard, Jean François

Abstract

The invention relates to a method for producing a turbomachine compressor blade, comprising the following steps:—installing primary pins (26) comprising a material other than a titanium-based alloy in primary bores (20) of a core, the primary bores forming at least one polygon, and installing a secondary pin made of titanium-based alloy in a secondary bore of the core; —producing a stack (2) of a suction-face sheet (4), a core (14) and a pressure-face sheet (6); —compacting the stack; —removing the primary pins (26) from the primary bores (20); —removing the secondary pin from the secondary bore; and—taking the core (14) away from the stack.

IPC Classes  ?

  • B23P 15/04 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from several pieces
  • B21D 53/78 - Making other particular articles turbine blades
  • B23K 20/02 - Non-electric welding by applying impact or other pressure, with or without the application of heat, e.g. cladding or plating by means of a press
  • F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion

85.

METHOD FOR SURFACE TREATMENT BY SELECTIVE REMOVAL OF A BONDING PRIMER ON A TITANIUM OR TITANIUM ALLOY SUBSTRATE

      
Application Number 18256184
Status Pending
Filing Date 2021-11-29
First Publication Date 2024-01-18
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Grunenwald, Anthony
  • Bossan, Pierre Antoine

Abstract

A method for treating the surface of a titanium or titanium alloy metal reinforcement of a blade made of composite material enables selective removal of a bonding primer with respect to the titanium or titanium alloy reinforcement. The method includes subjecting the metallic reinforcement to a thermal treatment performed at a temperature of between 250 and 350° C. for a period of between 1 hour and 10 hours in an oxidizing atmosphere. The method further includes subjecting the metallic reinforcement, after the thermal treatment, to a chemical pickling in an alkaline bath.

IPC Classes  ?

86.

TURBINE ENGINE FOR AN AIRCRAFT

      
Application Number 18256259
Status Pending
Filing Date 2021-12-02
First Publication Date 2024-01-18
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor De Carné-Carnavalet, Vincent Marie Jacques Rémi

Abstract

A turbine engine extends along an axis (X) and includes a flow path of a primary flow (F1) that has a compressor, a combustion chamber, and a turbine. The turbine engine further includes an outflow region of a secondary flow (F2) that surrounds the primary path, a blower or a propeller located upstream of the primary path and the outflow region of the secondary flow (F2), at least one arm extending radially through the primary path, and at least one fluid circulation pipe extending inside the arm. The arm includes an inlet for air from the primary path so as to cool the fluid circulating in the pipe.

IPC Classes  ?

  • F02C 7/14 - Cooling of plants of fluids in the plant
  • F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air

87.

CASE COMPRISING INTERNAL AND/OR EXTERNAL STIFFENERS

      
Application Number 18036035
Status Pending
Filing Date 2021-11-17
First Publication Date 2024-01-18
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Duval, Sylvain Yves Jean
  • Genilier, Amaud Lasantha
  • Renon, Olivier

Abstract

The invention relates to a turbine engine case (1), extending around an axis, said case (1) comprising: a hub comprising an outer wall (21) and an inner wall (22) between which an annular flange (23) extends; an outer shell (3); a plurality of arms (4) which each extend between the hub and the outer shell (3); the case (1) further comprising a plurality of main stiffeners (5), each being disposed in the extension of an arm (4) between the outer wall (21) and the inner wall (22) projecting from the flange (23), each main stiffener (5) comprising two ribs (51, 52) arranged opposite one another and which move away from one another from the outer wall (21) towards the inner wall (22).

IPC Classes  ?

  • F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
  • F01D 25/04 - Antivibration arrangements

88.

HEAT EXCHANGER MOUNTED IN A TURBINE ENGINE CAVITY

      
Application Number 18256200
Status Pending
Filing Date 2021-12-07
First Publication Date 2024-01-18
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Oriol, Sébastien
  • Cotereau, Nicolas Vincent Pierre-Yves
  • Boutaleb, Mohammed-Lamine

Abstract

A heat exchange system for a turbine engine is provided. The heat exchange system includes a cavity having an air intake, a heat exchanger arranged in the cavity and having a first circuit in which a first fluid can circulate, a movable flap mounted at the air intake and moving between two positions permitting or preventing, respectively, the circulation of air flow in the cavity, and a control device having a movable member configured to drive the movement of the movable flap. The control device can be arranged in the heat exchanger supply circuit and configured so as to permit or prevent the circulation of the first fluid to the heat exchanger and simultaneously move the movable flap between at least one of the two positions.

IPC Classes  ?

  • F02C 7/14 - Cooling of plants of fluids in the plant

89.

OIL RECOVERY DEVICE FOR A TURBOMACHINE

      
Application Number 18348870
Status Pending
Filing Date 2023-07-07
First Publication Date 2024-01-11
Owner
  • SAFRAN AIRCRAFT ENGINES (France)
  • SAFRAN TRANSMISSION SYSTEMS (France)
Inventor
  • Pennacino, Antoine Jacques Marie
  • Chassagne, Amélie Argie Antoinette
  • Gedin, Patrice Jocelyn Francis
  • Marcos Izquierdo, Juan-Luis
  • Peltier, Jordane Emile André

Abstract

An oil recovery device for a reducer of an aircraft turbomachine, the device having a gutter extending around an axis and being intended to be arranged facing oil ejection means formed in a ring gear of the reducer of the turbomachine. The gutter can include at least one oil reception chamber which extends around the axis and which is open towards the axis so as to receive the oil coming from the oil ejection means. The reception chamber can be delimited by a first wall and a second wall arranged facing each other. The first wall can have an axial oil evacuation orifice and the second wall can have a radial dimension relative to the axis which increases increasingly along the axis towards the axial orifice, so as to allow oil ejected by the reducer to be conveyed to the axial orifice.

IPC Classes  ?

  • F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
  • F02C 7/06 - Arrangement of bearings; Lubricating

90.

DEVICE FOR SETTING THE PITCH OF BLADES FOR A TURBINE ENGINE, AND TURBINE ENGINE COMPRISING SAME

      
Application Number 18546405
Status Pending
Filing Date 2022-03-07
First Publication Date 2024-01-11
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Oriol, Sebastien
  • Frantz, Caroline Marie
  • Verdier, Bastien Pierre

Abstract

A device for setting the pitch of blades for a turbine engine, including a control system acting on a linkage mechanism and including an actuator having two chambers; —a supply member coupled to a fluid transfer bearing for transferring a pressurised fluid from a main supply source in a fixed frame of reference to the chambers; —an auxiliary supply circuit which is arranged in a rotating frame of reference and is connected to the supply member and to the main supply source, the supply member being configured to: —in normal operation, allow fluid to pass from the main supply source to the chambers and to the auxiliary supply circuit, and —in the event of a break in the supply to the chambers from the main supply source, allow pressurised fluid to pass from the auxiliary supply circuit to the chambers.

IPC Classes  ?

  • F01D 7/00 - Rotors with blades adjustable in operation; Control thereof
  • F15B 13/04 - Fluid distribution or supply devices characterised by their adaptation to the control of servomotors for use with a single servomotor
  • F15B 1/04 - Accumulators
  • B64D 27/10 - Aircraft characterised by the type or position of power plant of gas-turbine type

91.

METHOD FOR ADDITIVE MANUFACTURING OF A WALL FOR A TURBINE ENGINE, COMPRISING AT LEAST ONE COOLING APERTURE

      
Application Number 17995203
Status Pending
Filing Date 2021-03-26
First Publication Date 2024-01-11
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Antony, Pierre
  • Pouzet, Sébastien Yohann

Abstract

A method is provided for additive manufacturing of a wall for a turbine engine, the wall including a first cooling aperture. The manufacturing method includes additively manufacturing the wall by selective melting or selective sintering on a powder bed. The wall is manufactured at least partially around the first cooling aperture with at least one lower zone that has a minimum length between 0.01 and 0.4 mm and/or with a lower zone that has a total thickness between 0.06 and 0.22 mm.

IPC Classes  ?

  • B22F 5/00 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
  • B22F 10/28 - Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM]
  • F23R 3/00 - Continuous combustion chambers using liquid or gaseous fuel

92.

BLADE COMPRISING A SHIELD HAVING A DEFROSTING AIR PASSAGE DUCT

      
Application Number 18037872
Status Pending
Filing Date 2021-11-10
First Publication Date 2024-01-04
Owner Safran Aircraft Engines (France)
Inventor
  • Gondre, Guillaume Pascal Jean-Charles
  • Leger, Morgane Astrid
  • Mahias, Stéphane Roger
  • Mauvais, Romuald Bernard

Abstract

The invention relates to a blade comprising a shield (14) attached to the upstream end of the body of the blade, the shield (14) comprising an upstream end forming a leading edge of the blade, the shield (14) further comprising a nose (141) upstream from which the leading edge is located, an pressure-face fin and an suction-face fin laterally attached on the blade, the fins extending from the nose, the blade comprising a defrosting air passage duct (15) arranged inside the nose and extending radially inside the nose, the duct (15) having an inner end (151) emerging opposite the root (125) and a radially external end emerging from the nose between the leading edge and its junction at the pressure-face fin.

IPC Classes  ?

  • F01D 25/02 - De-icing means for engines having icing phenomena
  • F01D 5/14 - Form or construction

93.

MODULARITY OF AN AIRCRAFT TURBOMACHINE

      
Application Number 18251379
Status Pending
Filing Date 2021-11-02
First Publication Date 2024-01-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Levisse, Paul Ghislain Albert
  • Belmonte, Olivier

Abstract

An aircraft turbomachine, including a first shaft having external splines oriented parallel to an axis of rotation of the shaft, a second shaft having internal splines complementary to the external splines, this second shaft being fitted axially on the first shaft and these shafts rotating as one by engagement of their splines, and a system for axially locking the shafts with respect to one another, wherein the system includes screws that are oriented radially with respect to the axis, each of these screws being screwed into a first orifice in one of the shafts and having a free end for engaging a second orifice in the other of the shafts.

IPC Classes  ?

  • F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
  • F16D 1/02 - Couplings for rigidly connecting two coaxial shafts or other movable machine elements for connecting two abutting shafts or the like

94.

COMPOSITE PART, IN PARTICULAR FOR AN AIRCRAFT TURBINE ENGINE

      
Application Number 18252753
Status Pending
Filing Date 2021-11-17
First Publication Date 2024-01-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Lhommeau, Tony Alain Roger Joël
  • Lamouche, Damien Bruno
  • Minervino, Mattéo
  • Lavigne, Julie Valérie Clara

Abstract

A part made of composite material, in particular for an aircraft turbine engine, includes a body produced from fibers woven in three dimensions and immersed in a resin, in particular an organic or polymer resin. The part has at least one crack sensor capable of detecting a crack in the part, at least one link element connected to the sensor and configured to transmit a signal (S), in particular to a member outside of the part, when a crack in the part is detected. The crack sensor has at least one conductive wire woven with the fibers of the body, and the link element is configured to measure the resistance of the conductive wire and to wirelessly transmit the signal.

IPC Classes  ?

  • F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
  • F01D 21/00 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for

95.

METHOD AND DEVICE FOR MANUFACTURING A DUAL-MATERIAL TURBINE ENGINE DISC AND DISC PRODUCED USING SAID METHOD

      
Application Number 18253510
Status Pending
Filing Date 2021-11-16
First Publication Date 2024-01-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor Longuet, Arnaud Cyril

Abstract

A method for manufacturing a dual-material turbine engine disc, includes the following operations: providing a rough bore made of a first material, mounting the rough bore about an axis of rotation of a rotating device, rotating the rough bore, spraying a second material under solidification conditions, thereby generating a column-like or monocrystalline microstructure, which is different from the first material, on an outer surface of the rough bore in order to produce a dual-material part, and machining the dual-material part to produce a turbine engine disc.

IPC Classes  ?

  • F01D 5/02 - Blade-carrying members, e.g. rotors
  • F01D 5/34 - Rotor-blade aggregates of unitary construction
  • C23C 4/12 - Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge characterised by the method of spraying
  • C23C 4/08 - Metallic material containing only metal elements
  • C23C 4/18 - After-treatment

96.

AIRCRAFT TURBINE ENGINE ASSEMBLY COMPRISING A PASSIVE VALVE FOR BY-PASSING A FUEL/OIL HEAT EXCHANGER

      
Application Number 18253791
Status Pending
Filing Date 2021-11-30
First Publication Date 2024-01-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Oriol, Sébastien
  • Cotereau, Nicolas Vincent Pierre-Yves
  • Boutaleb, Mohammed-Lamine

Abstract

An aircraft turbine engine assembly, including a lubricating oil circuit and a fuel supply device, the assembly also including a passive bypass valve enabling the oil to bypass an exchanger, the valve including a valve body and a piston arranged to move in a sliding space defining a first actuating chamber supplied with oil from the oil circuit, as well as a second actuating chamber supplied with fuel from the device, the valve being configured so that when the differential pressure between the oil pressure in the first actuating chamber and the fuel pressure in the second actuating chamber drops below a predetermined value of differential pressure, the piston moves from a normal operating position to an exchanger bypass position.

IPC Classes  ?

  • F02C 7/224 - Heating fuel before feeding to the burner
  • F02C 7/14 - Cooling of plants of fluids in the plant

97.

NACELLE AIR INTAKE FOR AN AIRCRAFT PROPULSION ASSEMBLY TO PROMOTE A THRUST REVERSAL PHASE

      
Application Number 18250914
Status Pending
Filing Date 2021-10-28
First Publication Date 2024-01-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Cheraga, Yacine
  • Mincu, Daniel-Ciprian
  • Chanez, Philippe Gérard
  • Lecordix, Jean-Loïc Hervé

Abstract

An air intake for an aircraft propulsion assembly extending along a longitudinal axis and comprising a turbine engine that comprises a primary flow path and a secondary flow path for respectively guiding a primary air flow and a secondary air flow during a thrust, and thrust reversal means for changing the secondary airflow into a reverse airflow during a thrust reversal. The air intake comprising a peripheral external enclosure comprising, in each plane radial to the axis, a point of maximum curvature for detaching the reverse airflow, an osculating circle defining a radius of curvature that is defined at each of the points of maximum curvature. The average value of the radii of curvature being less than a product of 0.028 times an internal radius of the intake at the fan.

IPC Classes  ?

  • B64C 7/02 - Nacelles
  • B64D 33/02 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
  • F02C 7/042 - Air intakes for gas-turbine plants or jet-propulsion plants having variable geometry

98.

IMPROVED TURBINE RING ASSEMBLY

      
Application Number 18251930
Status Pending
Filing Date 2021-11-04
First Publication Date 2024-01-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Cazin, Clément Emile André
  • Tabarin, Pascal Cédric
  • Congratel, Sébastien Serge Francis
  • Tableau, Nicolas Paul

Abstract

A turbine ring assembly extending around a longitudinal axis and comprises: a plurality of ring angular sectors made of CMC material circumferentially arranged in such a way as to form a turbine ring, each sector comprising a base from which an upstream leg and a downstream leg extend radially, a ring support structure comprising: a spacer extending around the sectors and comprising a downstream flange in such a way that the downstream leg of each sector is held against the downstream flange, a upstream force-absorbing plate extending circumferentially around the longitudinal axis in such a way as to be in contact with the upstream leg of each sector and fixed to an upstream area of the spacer which extends in extension of the upstream leg.

IPC Classes  ?

  • F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
  • F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector

99.

Turbojet engine rear part comprising a nozzle having flaps comprising levers that are movable by means of upstream and downstream bearing walls

      
Application Number 18252079
Grant Number 12018628
Status In Force
Filing Date 2021-11-05
First Publication Date 2024-01-04
Grant Date 2024-06-25
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Kohn, Thierry
  • Eichstadt, Frédéric Paul
  • Fremont, Elric Georges André

Abstract

A rear part for a turbojet engine includes a variable-geometry nozzle in which a convergent flap is provided with a lever supporting a bearing roller arranged axially between an upstream bearing wall and a downstream bearing wall, which are rigidly secured to a movable part capable of moving axially on command with respect to an upstream stator structure in such a way that the bearing roller is free to move with respect to the upstream and downstream bearing walls in a radial direction. During a downstream movement of the movable part, the upstream bearing wall pushes the lever to pivot a downstream end of the convergent flap towards the longitudinal axis. Operation of the drive mechanism for moving the movable part can thus be optimized while limiting the size and the mass of the mechanism for controlling the flaps.

IPC Classes  ?

  • F02K 1/12 - Varying effective area of jet pipe or nozzle by means of pivoted flaps
  • F02K 1/06 - Varying effective area of jet pipe or nozzle

100.

AIRCRAFT TURBINE ENGINE DRIVE LINE SHAFT COMPRISING A CONNECTING RING FOR DISASSEMBLING THE DRIVE LINE

      
Application Number 18252119
Status Pending
Filing Date 2021-11-08
First Publication Date 2024-01-04
Owner SAFRAN AIRCRAFT ENGINES (France)
Inventor
  • Jacquemard, Christophe Paul
  • Agneray, Xavier Jean Yves Alain

Abstract

Disclosed is a main shaft for an aircraft turbine engine drive line extending longitudinally from front to rear along an axis and having a rear portion configured for receiving a front portion of a rear shaft by interlocking, the main shaft having: —a front retaining member and a rear retaining member together defining a prison and each having a central opening, —a connecting ring, located in the prison, having a rear end configured to be screwed to a front end of the rear shaft extending into the central opening of the rear retaining member, the connecting ring having a gripping member formed on the inner surface and accessible from a front portion of the main shaft by a tool extending via the central opening of the front retaining member.

IPC Classes  ?

  • F01D 5/02 - Blade-carrying members, e.g. rotors
  • F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
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