Safran Aircraft Engines

Snecma

Thierry Georges Paul Papin

Matthieu Arnaud Gimat

Cedric Zaccardi

Kaelig Merwen Orieux

Regis Eugene Henri Servant

Christophe Marcel Lucien Perdrigeon

Dominique Raulin

Clément Cottet

Mohamed-Lamine Boutaleb

Vivien Mickaël Courtier

Vincent Joudon

Cédric Zaccardi

Sebastien Christophe Chalaud

Virginie Guilbert

Eric Pierre Georges Lemarechal

Christophe Paul Jacquemard

Dominique Marie Christian Coupe

Francois Marie Paul Marlin

Clement Cottet

Lise Lacroix

Thierry Francois Maurice Duchatelle

Régis Eugène Henri Servant

Jacky Novi Mardjono

Jean-Marc Claude Perrollaz

Yann Marchal

Alain Marc Lucien Bromann

Didier Gabriel Bertrand Desombre

Fabien Roger Gaston Caty

Alexandre Branco

Marc-Antoine Anatole Got

Cedric Magret

Arnaud Sanchez

Azath Mohamed

Xavier Jean Yves Alain Agneray

François Marie Paul Marlin

Fabrice Cretin

Damien Lourit

Alain Raffy

Mathieu Daniel Marie Van Gansen

Kaëlig Merwen Orieux

Yvan Guezel

Thierry François Maurice Duchatelle

Virginie Martine Collin

Sebastien Vincent Francois Dreano

Eric Pierre Georges Le Marechal

Cédric Magret

An assembly for an aircraft turbine engine is provided. The assembly generally includes an un-ducted propeller blade of which the root has a bulbous portion; a system for setting the pitch of the blade about a pitch setting axis, the pitch setting system including a cup configured to be placed inside the outer casing, has a flared open top end for axial insertion of the root and in which the root of the blade can be axially affixed by a locking member; and at least one radial retention prong in a radial space reserved between the wall of the cup and the root.

Assembly comprising a vane and a vane pitch setting system

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.

Sealing device for a discharge valve of a turbomachine

An assembly for an aircraft turbine engine, includes a fan, a speed reduction gear and an intermediate casing including an internal annular structure that houses an epicyclic gear train equipped with a planet carrier. The assembly includes a secondary retention device for the ring gear of the planet carrier, the device including: first projections that project radially inwards from the internal annular structure, each first projection having a recess that is open radially inwards, axially downwards and circumferentially in a first direction; and second projections that project radially outwards from the ring gear of the planet carrier and cooperate with the first projections in such a way that each second projection is partially accommodated in the recess of the first projection with which it is associated and downstream therefrom.

Assembly for an aircraft turbine engine, comprising means for the axial and radial retention of a fan

An assembly for an aircraft turbine engine is provided. The assembly generally includes an un-ducted propeller blade of which the root has a bulbous portion; a system for setting the pitch of the blade about a pitch setting axis, the pitch setting system including a cup, configured to be placed inside the outer casing, has a flared open top end for axial insertion of the root and in which the root of the blade can be axially affixed by a locking member; and at least one radial retention prong in a radial space reserved between the wall of the cup and the root.

The invention relates to a turbomachine module (), comprising: &#x2014;a rotating housing (-) supporting a propeller provided with a plurality of blades (), &#x2014;a system for varying the pitch of the propeller blades (), the system comprising a control means, and a mechanism for varying the pitch of the propeller blades, characterised in that the system is supported by the rotary housing (-), in that the control means comprise an annular row of rotary actuators (), and in that the mechanism for varying the pitch of the blades comprises a synchronisation ring () that is driven to rotate by rotary output shafts () of the actuators (), the synchronisation ring () being guided in rotation relative to the rotary housing (-) by guide means and meshed by a first toothing () with pinions () of the blades ().

Turbomachine module for a propeller having variable-pitch blades and turbomachine comprising same

The invention relates to a module of a turbomachine of longitudinal axis X, the module comprising: &#x22c5;&#x2014;a rotary casing () that rotates about the longitudinal axis and bears a propeller () which is provided with a plurality of blades (); &#x22c5;&#x2014;a fixed casing () comprising a cylindrical wall () extending between an inner wall () and an outer wall () of the rotary casing (); and, &#x22c5;&#x2014;a system () for changing the pitch of the blades of the propeller, mounted around the fixed casing and comprising: a control means () comprising a movable body () that is able to move axially on said fixed casing, at least one load transfer bearing () comprising an inner ring () that is connected to the movable body (), and an outer ring (); and a connection mechanism () for connecting the outer ring () to the blades of the propeller. According to the invention, the module further comprises a feathering device () for feathering the blades of the propeller, this device () comprising an annular row of springs () that are arranged around said fixed casing and extend axially, the springs () comprising first axial ends () that bear against a ferrule () of the fixed casing, and opposite second axial ends () that bear against said inner ring ().

Turbomachine module equipped with a system for changing the pitch of the blades of a propeller and with a device for feathering the blades

A turbomachine module with a longitudinal axis, the module comprising: a rotary casing rotatable around the longitudinal axis and arranged to carry a propeller provided with a plurality of blades; a system for changing the pitch of the blades of the propeller comprising: a control, and a mechanism for varying the pitch of the blades of the propeller connecting these blades to the control, wherein the control comprises a rotary actuator comprising a control body and a reference body which is integral with the rotary casing, and wherein the mechanism for varying the pitch comprises a synchronization ring which is driven in rotation around the longitudinal axis by the control body and which is guided in rotation on the rotary casing, the synchronization ring being connected to the blades.

Turbomachine module for a variable pitch blade propeller and turbomachine comprising it

The invention relates to an intermediate case () for a twin spool turbomachine for an aircraft, comprising a hub (), an outer shell () and outlet guide vanes () installed at their ends on the hub and on the outer shell, and each of at least some of the outlet guide vanes () performing a heat exchanger function and comprising a lubricant passage () designed to be cooled by the fan flow () following an outer surface of the outlet guide vane. According to the invention, the case also comprises at least one lubricant duct () passing along a circumferential direction of the hub () and at least part of which is made from a single casting with the hub, the lubricant duct () having at least one lateral opening communicating with the lubricant passage () of at least one of the vanes ().

Intermediate case for an aircraft turbomachine made from a single casting with a lubricant duct

A vaned ring for a turbomachine stator includes an outer shell, a vane having a vane head, and two platforms arranged circumferentially on either side of the vane head, and fixed on a radially inner face of the outer shell by detachable fasteners. The vane head has first conical seatings bearing on respective second conical seatings of the two platforms so as to prevent a displacement of the vane radially inwards and along the circumferential direction. The vaned ring includes a frangible pin including a base mounted on the outer shell and a head projecting radially inwards with respect to the outer shell and having a head surface forming, vis-&#xe0;-vis a top surface of the vane head, a stop opposing a displacement of the vane radially outwards.

Vaned ring for turbomachine stator having vanes connected to an outer shell by conical seating and frangible pin

A ring of vanes for an aircraft turbine engine, the ring presenting an axial direction and a radial direction and including a first annular sheath presenting an inside surface and a second annular sheath presenting an inside surface facing the first inside surface of the first annular sheath, the first and second sheaths being coaxial and defining between them a flow passage for a gas stream. The ring further includes both partially annular acoustic treatment panels including resonant cavities and also vanes extending in the radial direction between the first and second sheaths. Each vane includes a radial airfoil with at least two attachment tabs at each radial end of the airfoil fastened to the first or second sheath, and the inside surface of at least one of the first and second annular sheaths and the corresponding attachment tabs are covered by acoustic treatment panels arranged between the airfoils.

Ring of turbojet vanes including an acoustic treatment structure

A guide vane for a twin-spool aircraft turbomachine has an aerodynamic part that includes an internal lubricant cooling passage extending along a principal lubricant flow direction. The aerodynamic part is made in a single piece and also includes heat transfer fins arranged in the passage connecting the intrados and extrados walls and extending approximately parallel to the direction, these fins being distributed in successive rows along the principal direction and made such that for two rows of staggered directly consecutive fins, a first row includes fins forming a positive acute angle A1 with a dummy reference plane, while a second row includes fins forming a negative acute angle A2 with this plane.

Outlet guide vane for aircraft turbomachine, with improved lubricant cooling function

The invention relates to a guide vane intended to be mounted in a turbine engine between an inner shroud () and an outer shroud (), comprising a longitudinal straightening body () for an air flow extending between a first end () intended to be positioned at the inner shroud () and a second end () intended to be positioned at the outer shroud (), the longitudinal straightening body having an aerodynamic profile defined by a leading edge () and a trailing edge () in the flow direction of the air flow, and by a camber line () extending from the leading edge () to the trailing edge (). It further comprises a first attachment heel () and a second attachment heel () positioned in the continuation of the longitudinal body () at the first end () and the second end () respectively, the first () and second () attachment heels being planar elements arranged parallel with respect to one another, each attachment heel () being arranged at a distance from the leading edge () and from the trailing edge ().

Turbine engine flow guide vane with removable attachment

A bypass turbine engine is described. The engine includes a system for bleeding gas from a primary stream to a secondary stream. The system may include two coaxial rings extending one around the other. A perforated internal first ring may be situated at the level of an outer casing and may extend substantially in a continuation of the outer casing. A perforated external second ring may be mounted such that it slides circumferentially on the first ring between a first position in which the perforations of the rings do not communicate with one another and a second bleed position in which the perforations of the rings communicate with one another.

Bypass turbomachine fitted with bleed system

A guide vane for a dual flow aircraft turbine engine, the aerodynamic part of the vane including an inner duct for lubricant cooling extending in a main direction and being partly bounded by a pressure side wall and a suction side wall of the vane. A heat conduction matrix is lodged in the duct, and presents main heat transfer wings extending parallel to the direction, and laid out in staggered rows.

Output director vane for an aircraft turbine engine, with an improved lubricant cooling function using a heat conduction matrix housed in an inner duct of the vane

A turbomachine fan frame including a central hub, a plurality of radial vanes installed on the hub and distributed around the hub, each vane including an inner radial end root at which the vane is fixed to the hub, a plurality of flow stream reconstitution platforms, each platform being installed between the roots of two adjacent vanes and radially at a distance from the hub, wherein each platform is fixed to the hub by at least one support element associated with the platform, the support element located between the associated platform and the hub and located between the roots of two adjacent vanes.

Turbomachine fan frame comprising improved attachment means

Growing community of inventors

Varennes-Jarcy, France

Thierry Georges Paul Papin