Snecma

Safran

Safran Aircraft Engines

Matthieu Leyko

Jean Bertucchi

Pierre Jean-Baptiste Metge

Eric Conete

Christian Sylvain Vessot

Michel Pierre Cazalens

Julien Szydlowski

Gautier Mecuson

Mathieu Gaillot

Julien Antoine Henri Jean Szydlowski

Guillaume Bodard

Jonathan Langridge

Arthur Droit

Imane Ghazlane

A constant volume combustion system includes at least one combustion chamber having at least one admission port and an exhaust port. The system also includes at least one elastically deformable tongue made of ceramic matrix composite material forming an air admission valve, the tongue being present inside the chamber and being positioned facing the admission port, the tongue having a first end that is stationary relative to an inside wall of the chamber and a second end, opposite from the first end, the second end being free and movable relative to the inside wall.

Constant volume combustion system

A constant-volume combustion system for a turbomachine includes a plurality of combustion chambers distributed in an annular manner about an axis defining an axial direction, each combustion chamber including an intake port and an exhaust port; a selective closure member rotationally movable about the axis with respect to the combustion chambers, the selective closure member including a ferrule facing the intake and exhaust ports of the combustion chambers, the ferrule containing at least one intake aperture intended to cooperate with the exhaust port of each chamber and at least one exhaust aperture intended to cooperate with the exhaust port of each chamber. Each intake aperture and each exhaust aperture are segmented by at least one segment extending in each aperture in the axial direction.

Constant-volume combustion system comprising a rotating closure element with segmented apertures

The invention relates to a turbine engine combustion module (), in particular for an aircraft turbine engine, designed to carry out constant-volume combustion, comprising: at least two combustion chambers (A,B) arranged about an axis, each chamber (A,B,C) comprising a compressed gas intake port () and a burnt gas exhaust port (); and an ignition means that triggers combustion in the combustion chambers (A,B,C). The module () comprises at least one duct () which establishes a communication between a first combustion chamber (A) and at least one second combustion chamber (B) in order to inject burnt gases from the first combustion chamber (A) into the second combustion chamber (B) so as to trigger combustion in the second combustion chamber (B).

Constant-volume combustion module for a turbine engine, comprising communication-based ignition

A constant volume combustion chamber for a turbine engine, includes an intake port, an exhaust port, and a first rotary shutter facing the intake and exhaust ports and configured to rotate around an axis in a first direction of rotation, the first shutter including an aperture intended to cooperate alternately with the intake and exhaust ports during the rotation of the first shutter. The chamber further includes at least one second rotary shutter facing the intake and exhaust ports and configured to rotate around the axis in a second direction of rotation opposite to the first direction, the second shutter including an aperture intended to cooperate alternately with the intake and exhaust ports during the rotation of the second shutter, the first and second shutters being synchronized and configured so that their respective apertures intersect alternately when both are facing the intake and when both are facing exhaust ports.

Constant volume combustion chamber with counter rotating shutter valve

The disclosure relates to a flow deflector for a discharge valve system of a double flow turbomachine compressor. The flow deflector comprises a wall provided with a plurality of ejection channels positioned to discharge a discharge air flow. The ejection channels are arranged in rows along aligned directions, substantially parallel to an axial plane of the deflector. The ejection channels of each row are oriented at a defined angle between a normal line to the wall and axes of the ejection channels, and decreasing between an upstream edge and a downstream edge of the wall, as defined according to the direction of movement of the air flow.

Flow deflector for a discharge valve system, discharge valve system and turbomachine comprising such a discharge valve system

Heat-exchange and noise-reduction panel for a propulsion assembly, in particular for an aircraft, the panel comprising: a perforated plate comprising a plurality of through-openings; a cellular structure comprising longitudinally oriented structural walls covered by said perforated plate and comprising, between said walls, cavities that define Helmholtz resonators, said through-openings forming necks of said resonators; and means for the circulation of fluid, for example oil, at said perforated plate, wherein said fluid circulation means comprise channels that are formed at least in part in thickened ends of said walls on the same side as said perforated plate, and/or at least in part in regions of the perforated plate situated in the longitudinal extension of said thickened ends.

Heat-exchange and noise-reduction panel for a propulsion assembly

A combustion module of a turbine engine, in particular of an aircraft, is configured for carrying out constant-volume combustion. The module includes a plurality of combustion chambers angularly distributed in a regular manner around an axis. Each chamber has an intake port for pressurized gas and an exhaust port for combustion gases. Each intake/exhaust port is configured to be opened or closed by a corresponding common rotating intake/exhaust valve which is coaxial with the axis.

Constant-volume combustion module for a turbine engine

A discharge valve system of a bypass turbine engine compressor includes a flow deflector. The flow deflector has a wall provided with a plurality of ejection channels configured to discharge a discharge airflow from the compressor in a duct of the turbine engine in which an airflow circulates. The ejection channels are disposed in a plurality of concentric rows around a center and are oriented at an angle defined between a straight line normal to the wall and the axes of the ejection channels. The angle of the ejection channels vary by increasing from the peripheral edge towards the center of the wall so that the discharge airflow passing through the ejection channels is guided in a gyratory movement around the center.

Gyratory-effect flow deflector of a discharge valve system, discharge valve system and turbine engine comprising such a discharge valve system

A forward fan rotor is disclosed with a hub of axis of rotation (X) and a cone mounted on the hub of the fan. The cone comprises an air bleed orifice which opens into an air duct of which a forward end portion passes through the fan rotor, said forward end portion comprising mechanical air entrainment means. The air bleed orifice has an annular shape and in that the cone is divided by said orifice into a front vertex portion and a rear frustoconical portion. A turbomachine forward axial spool equipped with such a fan rotor is also disclosed.

Fan rotor for a turbo machine such as a multiple flow turbojet engine driven by a reduction gear

A device for circulation of air for a turbomachine, the device including an inlet provided to receive incident airflow, air circulation ducts connected to the inlet, a central part including side walls, peripheral walls arranged at a distance from the central part, so as to form the air circulation ducts, each duct being arranged between one of the side walls of the central part and one of the peripheral walls, each peripheral wall including an upstream end portion delimiting with one of the side walls an inlet orifice of one of the ducts and a downstream end portion delimiting with one of the side walls an outlet orifice of one of the ducts, at least one member mobile in translation relative to the peripheral walls, the at least one mobile member being formed in the central part and arranged so that its displacement varies a section of each of the inlet orifices and a section of at least one of the outlet orifices.

Air circulation device for turbomachine

A discharge system of a separated twin-flow turbojet for an aircraft, supported by a suspension mast, is disclosed. The system includes a main nozzle delimited by an annular cowl with a slot of annular shape defining upstream and downstream portions of the cowl and which is traversed by the suspension mast. The downstream portion of the cowl of the main nozzle includes a first part extending downstream from the upstream portion of the cowl to a trailing edge of the main nozzle, on either side of the suspension mast along two predefined angular sectors; and a second part formed from an internal contour of the slot and having a trailing edge with a diameter smaller than that of the trailing edge associated with the first part of the downstream portion of the cowl. Connecting walls laterally connect the first and second parts of the downstream portion.

Discharge system of a separated twin-flow turbojet for an aircraft, corresponding turbojet and associated design method

A turbine engine convergent-divergent nozzle including an annular central element and an annular cap arranged coaxially around the central element to co-operate with the central element to define an annular flow channel for a gas stream from the engine. Between a throat section and an ejection section of the nozzle, the central element and the cap present respective internal profiles in longitudinal section that are modeled by curves having respective radii of curvature that are identical in absolute value.

Convergent-divergent nozzle for a turbine engine

A turbomachine including a nozzle extending along an axial direction and configured to convey a hot gas stream; and a mounting pylon for mounting the turbomachme, in which the pylon extends radially from a base arranged in the vicinity of the nozzle, is provided. The turbomachine includes a pylon cooler device having at least one auxiliary gas duct having a gas inlet configured to take cold gas from the outside of the nozzle, and a gas outlet opening out in the vicinity of the base of the pylon.

Turbomachine comprising a device for the cooling of a pylon

In an assembly including a bypass turbine engine and an attachment system attaching the turbine engine to an aircraft structure, the attachment system includes a mast, at least two rear attachments connecting the mast to two rear attachment points of the turbine engine, and a rear covering covering the rear attachments and providing a streamlined fairing. The rear covering includes two side walls, each including a first portion extending along and in proximity of the mast and extending rearwards, and a second portion projecting sideways relative to the first portion towards its outside, to cover the rear attachments. Each second portion is substantially a bulge and is connected to the first portion along a curved connection line extending between front and rear points situated substantially level with an inner longitudinal edge of the first portion to minimize the impact on a flow of bypass stream leaving the bypass passage.

Assembly formed by a turbine engine and a system for attaching it to an aircraft structure

According to the invention, the method for defining the shape of a convergent-divergent nozzle of longitudinal axis L-L comprising a convergent part connected, at a throat, to a divergent part the downstream free end of which comprises scallops delimiting noise-reducing chevrons distributed in the circumferential direction, comprises defining beforehand the divergent part of the nozzle using a divergent first portion, a divergent second portion and a convergent portion, attributing an initial value to dimensional parameters defining the chevrons and the three portions, calculating performance criteria and checking that the calculated performance criteria satisfy predefined performance conditions.

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Moissy-Cramayel, France

Matthieu Leyko