Safran Aircraft Engines

Snecma

Cédric Zaccardi

Christophe Marcel Lucien Perdrigeon

William Henri Joseph Riera

Christophe Paul Jacquemard

Thierry Georges Paul Papin

Laurent Jablonski

Jacky Novi Mardjono

Antoine Robert Alain Brunet

Catherine Pikovsky

Amélie Argie Antoinette Chassagne

Kaëlig Merwen Orieux

Guillaume Claude Robert Belmon

Sébastien Vincent François Dreano

Mohamed-Lamine Boutaleb

Marc-Antoine Anatole Got

Xavier Jean Yves Alain Agneray

François Marie Paul Marlin

Mathieu Daniel Marie Van Gansen

Mathilde Anaëlle Mélissia Saurat

Eric Pierre Georges Le Marechal

Gwenael Sébastien François Marion

A heat exchanger between a fluid and an air flow, includes a heat exchange wall separating the fluid and the air flow, the heat exchange wall including a heat exchange surface that extends parallel to a longitudinal direction of the air flow and with which the air flow is in contact. The heat exchange wall includes at least one turbulence generator extending in a hollow manner in relation to the heat exchange surface.

Heat exchanger comprising a baffle wall with hollow turbulence generators

A bypass turbomachine is disclosed. The turbomachine includes a longitudinal axis having an upstream blower impeller and a downstream air-flow straightening assembly of a secondary annular passage delimited radially on the inside by a radially inner shroud and radially on the outside by a radially outer shroud, blades extending between the radially inner and outer shrouds and being attached at a first end portion to the radially inner shroud and at a second end portion to the radially outer shroud, the blades including a useful portion extending between the first and second end portions and defining a lower face and an upper face. For each blade, in a plane perpendicular to the longitudinal axis, each aerodynamically useful portion is bent into a C-shape in the circumferential direction.

Air flow straightening stage for a turbomachine

A turbofan engine including an exhaust casing traversed by a primary flow and surrounded by a secondary flow and an outlet cone carried by this exhaust casing, the exhaust casing and the outlet cone together defining an internal space. The exhaust casing includes a hollow radial arm traversing the primary flow in order to convey part of the secondary flow so as to form a cooling flow for supplying the internal space, and the outlet cone is terminated by an opening for discharging the cooling flow.

Turbofan engine comprising an outlet cone cooled by its secondary flow

The invention relates to a straightening assembly () comprising two radially inner and outer coaxial shells () between which extend vanes () made of composite material, fixed to a first end portion on the radially inner shell and to a second end portion on the radially outer shell, characterized in that:

Turbine engine comprising a straightening assembly

An assembly for a turbomachine through which an air flow passes, includes a stator including guide vanes extending radially in relation to a longitudinal axis, at least one inter-vane platform extending between the radially outer ends of two circumferentially consecutive guide vanes, each inter-vane platform including an outer surface that faces the axis, a heat exchanger located downstream of the stator in relation to a direction of circulation of the air flow in the turbomachine during operation, this stator including a heat exchange surface extending in the extension of an inter-vane platform. At least one inter-vane platform located in the upstream extension of the heat exchange surface is provided with at least one turbulence generator on its outer surface.

Heat exchanger comprising an inter-vane wall provided with hollow turbulence generators

An aircraft turbine engine having a primary air flow path with low-pressure and high-pressure compressors, a secondary air flow path which is located around the primary path and runs coaxially thereto, the turbine engine including vanes distributed about a main axis of the turbine engine. A pressurized air circuit draws air between the low-pressure compressor and the high-pressure compressor or in the high-pressure compressor and supplies at least one component located close to a main axis of the turbine engine. The pressurized air circuit includes a heat exchanger between the stream of pressurized air and the stream of air flowing in the secondary path, the heat exchanger being arranged in at least one of the straightening vanes, where a heat exchanger pipe is arranged, the pipe having a pressurized-air inlet and a pressurized-air outlet that are located at the same radial end of the vane.

Turbine engine comprising a heat exchanger in the secondary path

A turbomachine blade including a body that extends mainly in a plane defined by a main axis and a longitudinal direction, which is defined by a lower surface wall, an upper surface wall, a leading edge located at a first longitudinal end of the body and a trailing edge located at a second longitudinal end of the body, wherein the body of the blade includes a plurality of first pipes that extend mainly along the direction of the main axis, for circulation of a gas flow, and a plurality of second pipes that extend mainly along the longitudinal direction, for circulation of a second gas flow.

Turbomachine comprising a heat exchanger in the secondary path

A turbojet engine including a shaft surrounded by a low-pressure rotor surrounded by a coaxial and independent high-pressure spool, this turbojet engine including from upstream to downstream: a fan driven by the shaft; a low-pressure compressor carried by the rotor; an inter-compressor casing; a high-pressure compressor and a high-pressure turbine carried by the high-pressure spool; an inter-turbine casing; a low-pressure turbine carried by the rotor; an exhaust casing; this turbojet engine including an upstream rotor bearing carried by the inter-compressor casing; a downstream rotor bearing carried by the exhaust casing; a gearbox downstream of the downstream bearing and through which the rotor drives the shaft; a downstream shaft bearing downstream of the downstream rotor bearing.

Double-flow turbojet engine assembly with epicycloidal or planetary gearbox

The invention relates to a straightening assembly for an air flow of a turbomachine, comprising two respectively coaxial radially inner and outer shrouds (), between which extend the vanes () made of composite material attached, at a first end portion () thereof, to the radially inner shroud and at a second end portion () thereof, to the radially outer shroud (), with the vanes () comprising a useful portion extending between the two end portions (). According to the invention, for each vane () in a plane perpendicular to the axis of the radially inner and outer shrouds (), a straight line passing through a junction between said first and second end portions () and the useful portion forms, with a radius of the radially inner shroud passing through the junction between said first end portion and the useful portion of said vane (), an angle &#x3b1; such that 0&#xb0;&#x3c;&#x3b1;&#x2264;90&#xb0;.

Air flow straightening assembly and turbomachine including such an assembly

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

The invention relates to a guide vane for a bypass aircraft turbomachine, its aerodynamic part comprising a first lubricant cooling interior passage in which heat transfer structures are arranged and a second lubricant cooling interior passage in which heat transfer structures are arranged, the aerodynamic part comprising a bent area connecting a lubricant output end of the first interior passage to a lubricant input end of the second passage, the bent area extending along a curved generatrix and being partly delimited by the intrados wall and the extrados wall of the vane. According to the invention, the bent area comprises one or more lubricant guide(s) arranged between the intrados and extrados walls of the vane, and each extending substantially parallel to the curved generatrix of the bent area.

Aircraft turbo machine exit guide vane comprising a bent lubricant passage of improved design

A gas turbine casing made of composite material, the casing presenting an axisymmetric shape and including, in an axial direction: an upstream flange; a retention zone presenting thickness greater than the remainder of the casing; and a downstream flange. The casing presents, in section on a plane defined by the axial direction and by a radial direction, a profile including at least first and second inclined portions that are inclined relative to each other and relative to the axis of the casing. The first and second inclined portions meet at a point situated between the retention zone and one of the flanges, this point corresponding to the vertex of an angle formed by the first and second inclined portions. The distance between the vertex of the angle and the mean radius of the casing lies in the range 1.5% and 4% of the mean radius of the casing.

Composite material casing of stiffening shape

The invention relates to an intermediate case for a twin spool turbomachine for an aircraft, comprising at least one lubricant passage and piece () associated with at least one of its vanes (), comprising a lubricant passage () extending along a passage axis () and opening up into the lubricant duct (). The passage axis () and the orifice axis () are also at a spacing from each other along a circumferential direction (), and a connection piece () through which a connection conduit () passes is firstly partially housed in the vane root () so as to create a communication between the connection duct () and the lubricant passage orifice (), and secondly is partially housed in an end-piece support device () or in the lubricant passage end-piece (), such that the connection conduit () communicates with the lubricant passage () in the end-piece.

Intermediate case for an aircraft turbomachine comprising a lubricant passage end-piece connected to a case vane by a connection piece

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

An outlet guide vane wheel includes guide vanes made of polymer matrix composite material reinforced by fibers, each having a vane root and a vane tip. The vane roots are fastened on a hub of the wheel by a first connection, and the vane tips are fastened on an outer shroud of the wheel by a second connection. The first connection includes a bearing plane secured to the hub and a first backing plate for securing to the hub, with the vane roots being sandwiched between the bearing plane and the first backing plate. The second connection includes a second backing plate for securing to the shroud, with the vane tips being sandwiched between the shroud and the second backing plate.

Growing community of inventors

Saint Maur des Fosses, France

Cédric Zaccardi