Safran Aircraft Engines

Safran

Safran Helicopter Engines

Safran Ceramics

Nicolas Paul Tableau

Lucien Henri Jacques Quennehen

Sébastien Serge Francis Congratel

Clément Jean Pierre Duffau

Antoine Claude Michel Etienne Danis

Clément Jarrossay

Matthieu Arnaud Gimat

Aurélien Gaillard

Gilles Gérard Claude Lepretre

Hubert Jean-Yves Illand

Pascal Cédric Tabarin

Clément Emile André Cazin

Maurice Guy Judet

Guillaume Pascal Jean-Charles Gondre

Fabrice Marcel Noel Garin

Dany Abboud

Julien Paul Schneider-Die-Gross

Mohammed El Badaoui

Denis Daniel Jean Boisseleau

Pascal Cedric Tabarin

Fabrice Marcel Noël Garin

Clement Marie Benoit Roussille

Clement Jean Pierre Duffau

Sebastien Serge Francis Congratel

Turbine shroud assembly comprising sections () made from CMC and forming a shroud () and a support structure (), each section having a base () with a radially internal face () and a radially external face (), from which there extend in a projecting manner an upstream attachment lug () and a downstream attachment lug (), the support structure comprising a collar (), from which there radially extend in a projecting manner towards the shroud an upstream radial flange () and a downstream radial flange (), by which the lugs of each section of the shroud are retained, the shroud () being retained by axial pins () which cooperate, on the one hand, with the upstream radial flange, via first and second annular end plates (), and directly with the downstream radial flange and, on the other hand, with the upstream and downstream attachment lugs, respectively.

Turbine shroud assembly

The invention relates to a fan blade () of a turbomachine () comprising a structure () made of composite material having, within the root () and/or the stilt () of the blade, a first layer () comprising the lower face () and having a first thickness (E) of between 10% and 25% of the total thickness (E), a second layer () comprising the suction face () and having a second thickness (E) of between 10% and 25% of the total thickness (E), and a central layer () extending between the first layer () and the second layer (), the shortening of the warp strands () in the first and second layers () being greater than the shortening of the warp strands () in the central layer ().

Differentiated shortening of strands of the fibrous reinforcement of a fan blade

A turbine ring assembly extending about an axis, includes 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 lug and a downstream lug spaced axially from each other extend radially outwardly, the ring support structure including a spacer having a clamp against which the downstream lug of the ring sectors is held, a first upstream flange against which the upstream lug of the ring sectors is held, and upstream of the first upstream flange, a second upstream flange for absorbing forces against which the first upstream flange is held.

Turbine ring assembly mounted on a cross-member

A stator wheel for a turbomachine turbine. This wheel includes an inner shroud formed by an upstream portion and a downstream portion configured to be assembled with the blades by axial translation in a respective direction.

Turbomachine stator assembly comprising an inner shroud having upstream and downstream portions assembled by axial translation

Method for monitoring a turbomachine comprising a stator and a rotor, device, system, aircraft and computer program product. The method comprises a step of acquiring an input signal which represents a deformation of the stator or rotor of the turbomachine. The input signal has been acquired by a deformation gauge which is attached to the stator or rotor. The input signal comprises a first component which represents deformations of the stator or rotor which are caused by rotation of the rotor vanes in relation to the stator, and a second component which represents deformations which are caused by different elements of the rotor vanes. The method comprises a step of re-sampling the input signal in order to obtain a re-sampled input signal comprising a predefined whole number of samples per revolution of the turbomachine rotor. The method also comprises a step of processing the re-sampled input signal, comprising: filtering the re-sampled input signal, making it possible to acquire a filtered input signal in which the second component is attenuated, and separating the filtered input signal into a plurality of third components, each third component representing a contribution to the deformation caused by a respective associated vane. The method also comprises a step of detecting, in a third component, contact between the vane associated with the third component and the stator.

Method for monitoring a turbomachine, device, system, aircraft and computer program product

A turbine ring assembly extending around a longitudinal axis and comprises:

Turbine ring assembly

Turbine comprising a casing and a nozzle including a metal outer shroud integral with the casing, a metal inner shroud, and a plurality of nozzle sectors of CMC forming a ring extending between the metal outer shroud and the metal inner shroud, each sector including a mast, an inner platform, an outer platform and at least one airfoil having a hollow profile defining an inner recess, each of the inner and outer platforms having an opening communicating with said inner recess, and the mast passing through said openings and the inner recess and being attached to said casing and connected with said nozzle sector. The mast comprises at least one fastening projection having at least one portion extending from a radially outer end of the mast in a direction opposite to the center of the mast and in a plane orthogonal to the radial direction.

Turbomachine turbine having CMC nozzle with load spreading

A turbomachine turbine nozzle extending around a central axis, including at least one radially outer shroud, at least one radially inner shroud, and at least one blade made of ceramic matrix composite material, distinct from the radially inner shroud and from the radially outer shroud, and extending radially between the radially inner shroud and the radially outer shroud, the blade being hollow and including a cavity opening at a radially inner end and at a radially outer end of the blade, the nozzle including at least one tubular mast arranged in the cavity of the blade and allowing routing the ventilation air passing through the cavity of the blade, the mast including a radially outer end attached to the radially outer shroud, and a radially inner end cooperating with a radial flange for positioning the radially inner shroud.

CMC guide vane for a turbomachine turbine

A turbine assembly () comprising: —a plurality of turbine ring sectors () made of ceramic-matrix composite material, —a ring support structure (), comprising an annular shroud (), and in addition —a plurality of angular spacer sectors () together forming an annular spacer (), said annular spacer () being, on the one hand, fixed to the turbine ring () and, on the other hand, fixed to said annular shroud (), characterized in that said turbine assembly () comprises at least one air diffuser (), which is configured to diffuse cooling air onto the radially outer face () of at least one of said turbine ring sectors (), and in that said at least one air diffuser () is mounted by being nested on one of said angular spacer sectors (), in a nested position.

Turbine assembly, and gas turbine engine provided with such an assembly

A turbine comprises a casing, an outer metal shroud, an inner metal shroud and an annular distributor having a plurality of CMC ring sectors, each sector comprising a mast, an inner platform, an outer platform and at least one blade having a hollow profile that defines an inner housing, the inner and outer platforms each having an opening communicating with said inner housing, and the mast passing through said openings and the inner housing and being secured to said casing and connected to said annular sector. Each blade comprises at least one first radial shoulder projecting axially towards the inside of the blade, and each mast comprises at least one second shoulder projecting axially towards the outside of the mast configured to radially cooperate with a first shoulder and radially press the blade against the mast.

Turbomachine turbine having a CMC nozzle with load spreading

A device for cooling a turbine shroud comprising at least one annular flange configured to be attached to an annular radial collar of a shroud support structure being arranged upstream, with respect to a circulation direction of an air flow, of the turbine shroud, and comprising at least one cooling air circulation channel, a diffuser configured to be attached to said annular radial collar downstream of the annular flange and comprising at least one intake channel in fluid communication with the circulation channel of the annular flange, and comprising an injection cavity comprising a plurality of injection holes and being configured to inject on a radially external face of the shroud, via the injection holes, the cooling air originating in the intake channel, and a particle filter arranged on an inlet section of the circulation channel of the annular flange, the particle filter comprising a plurality of openings.

Aircraft turbine shroud cooling device

An assembly for a turbomachine turbine includes at least one ring sector made of CMC material and a support casing including an upstream flange and a downstream flange between which each ring sector is disposed, each ring sector including a base that has a radially external face from which radially extend two lugs, the lugs of each ring sector being retained between the two flanges of the support casing by axial pins each engaged in one of the flanges of the ring support casing and in the lug of the ring sector facing said flange, the assembly further including, for each ring sector, at least one radial retaining pin screwed into the support casing and coming to radially bear against a lug of the at least one ring sector to retain it in position, and anti-rotation system for rotationally locking the radial retaining pin.

Assembly for a turbomachine turbine and associated turbomachine

A method for assembling/disassembling a turbine ring assembly including a turbine ring of ceramic matrix composite material and a ring support structure of metal equipped with a central shroud and with at least one single-piece annular plate positioned upstream of the ring with respect to the direction of an air flow passing through the assembly, the plate having in the radial direction a first free end and a second end opposite to the first end, the second end being removably attached to the first radial flange and shrink-fit to the central shroud, is disclosed. The method for assembling/disassembling includes, before the assembling/disassembling of the plate, a step of deformation of the central shroud and/or of the at least one annular plate.

Method for assembling and disassembling a turbine ring assembly

The invention relates to an assembly comprising a turbomachine casing () and a nozzle made of ceramic matrix composite material having a blade (), the frontside and backside surfaces of which delimit an internal cavity, and which is connected at its radially external end by a connecting part () of the nozzle to the casing (), said connecting part () extending substantially radially outwards, the blade () and said connecting part () being formed in one piece, said assembly being characterized in that the connecting part () is fastened to a first radial wall () integral with the casing () by fastening means ().

Attachment of a nozzle to a casing of a turbomachine

A turbine shroud sector made of ceramic matrix composite, of longitudinal axis X and which includes a base with a radially internal face, a radially external face from which there extend upstream and downstream tabs for attachment to a shroud support structure. The base includes a first slot and a second slot, which is arranged radially on the outside of the first slot. The slots are formed in the lateral face. A first and a second sealing strip rest against a radially internal wall of the first and second slots. The first strip has the overall shape of an omega and the shroud sector exhibits a first clearance which is defined radially between a central part of the first strip and the radially internal wall of the first slot, and a second clearance which is defined radially between this central part and a radially external wall of the first slot.

Growing community of inventors

Moissy-Cramayel, France

Nicolas Paul Tableau