Snecma

Safran Aircraft Engines

Laurent Gilles Dezouche

Jean-Pierre Lombard

Yvon Cloarec

Francois Maurice Garcin

Stéphane Roger Mahias

Alain Chatel

Patrick Edmond Kapala

Olivier Chantoiseau

Olivier Abgrall

Sophie Martine Jobez

Philippe Joly

Alain Jean Charles Chatel

Julien Marc Antoine Prevost

Jean-Pierre Francois Lombard

Samir Zaidi

A turbine stator sector includes a plurality of vanes extending along a radial direction between a first end and a second end and along an axial direction between a leading edge and a trailing edge. The sector further includes an internal shroud linked to the first end of the vanes and an external shroud linked to the second end of the vanes. The sector includes at least one annular portion forming all or part of the internal shroud or of the external shroud. The annular portion includes a first partition present at the junction with the first or the second end of the vanes and a second partition held spaced from the first partition along the radial direction by a three-dimensional structure including a plurality of cutouts.

Turbomachine stator sector having flexible regions subjected to high stress

The invention relates to a turbomachine blade comprising a vane. The vane comprises a leading edge, a trailing edge, a pressure face and a suction face spaced apart from each other and connecting the leading edge to the trailing edge. The vane further comprises at least one internal cavity between the pressure face and the suction face, wherein air is intended to flow. The vane comprises intersecting transverse partitions, downstream of the cavity, so that air can flow from the cavity in the direction of the trailing edge between the intersecting partitions.

Turbomachine blade, comprising intersecting partitions for circulation of air in the direction of the trailing edge

An aircraft engine annular shroud includes, on its internal surface, a peripheral annular groove able to accept a plurality of blade roots; an insertion opening for inserting the blade roots into the annular groove; and a closure device for closing off the insertion opening.

Aircraft engine annular shroud comprising an opening for the insertion of blades

An angular diffuser sector for a turbine engine compressor is provided. The sector includes an inner shroud; an outer shroud; a vane; a casing element mounted at least in part around one of the shrouds and having an open groove facing an axial end of the corresponding shroud; a wedge-forming fitting interposed between the casing element and the corresponding shroud, the wedge presenting a first surface that contacts with a bearing surface of the axial end of the corresponding shroud, and a second surface that is inclined relative to the first surface and that contacts with a corresponding sloping surface of the groove in the casing element; and a device that exerts an axial force on the wedge so as to keep its first surface in contact with the bearing surface and its second surface in contact with the sloping surface of the groove in the casing element.

Angular diffuser sector for a turbine engine compressor, with a vibration damper wedge

A turbomachine stator stage including two coaxial shrouds, respectively an inner shroud and an outer shroud with substantially radial vanes extending therebetween, the radially outer ends of the vanes being welded to the outer shroud and the radially inner ends of the vanes passing through orifices in an outer ring of the inner shroud, the inner ends being held by elastomer spacers in an annular cavity formed between the outer ring and an inner ring that is fastened to the outer ring.

Stator stage for turbomachine compressor

A sector of a compressor guide vanes assembly or a sector of a turbomachine nozzle assembly. It comprises an inner ring sector (), an outer ring sector () and a multitude of blades () connecting the inner ring sector () to the outer ring sector (). The outer ring sector () or the inner ring sector () comprises radial cuts () situated between two consecutive blades () in such a way as to split them into the same number of elementary sectors as there are blades. Housings () are provided, secant to the radial cuts (), damping inserts () being positioned in said housings.

Sector of a compressor guide vanes assembly or a sector of a turbomachine nozzle assembly

The balancing flyweight of a turbomachine rotor includes two pyramid shaped end parts each one having a base and an apex, and an intermediate part which connects the two bases of the end parts together. The two apexes are aligned on a longitudinal axis. The two end parts and the intermediate part exhibit, in cross section through a plane perpendicular to said the longitudinal axis, cross-sections having polygonal shapes centered on said longitudinal axis.

Balancing flyweight, rotor disk equipped therewith, rotor and aircraft engine comprising them

A device for rotationally blocking a retaining ring on a rotor disk three successive blocking hooks of the disk (), two cleats of the retaining ring, arranged on either side of a split in the retaining ring. The position of the cleats on the retaining ring is such that, when the retaining ring is in place in a groove situated in the rotor disk, the two cleats are in respective abutment against the first and third blocking hooks, and the split is covered by the second blocking hook.

Device for blocking a ring for axially retaining a blade, associated rotor disk and retaining ring, and rotor and aircraft engine comprising them

A rotor disk for a turbomachine, the disk carrying blades whose roots are retained in a peripheral groove of the disk, the groove including a window for inserting blade roots and a locking member that is a single elastically deformable piece and dimensioned to be received and retained in a notch of the groove while in its free state, and to be capable of being displaced to slide inside the groove while in a stressed and elastically deformed state.

Growing community of inventors

Le Coudray Montceau, France

Laurent Gilles Dezouche