Safran Aircraft Engines

Snecma

Sébastien Claude Cochon

Pierre-Hugues Ambroise Maxime Victor Retiveau

Vincent Paul Gabriel Perrot

Pierre-Alain Francis Claude Sebrecht

Damien Bernard Emeric Guegan

Mickael Cavarec

Arnaud Langlois

A variable-pitch stator blade includes an airfoil having a central portion with a first chord and a first skeleton line delimited by a leading edge and a trailing edge. An end portion has a second chord and a second skeleton line delimited by the leading edge and a downstream limit. A skeleton angle at a first length Iof the first chord is defined by a function G(), and the skeleton angle at a second length lof the second rope being chord is defined by a function G(). The absolute value of the average increase Aof G′() between the leading edge and the downstream limit is greater than the absolute value of the average increase Al of Gl′() between the leading edge and a point P, wherein the first length Icorresponds to the total length of the second chord.

Variable-pitch stator vane for an aircraft turbine engine

A vane stage with a longitudinal axis designed to be fitted in a turbine engine compressor. The van stage has an annular row of mobile vanes arranged upstream from an annular row of stator vanes. The annular row of stator vanes has a radially internal annular platform bearing radial blades, an upstream annular portion of which is arranged upstream from the blades and is surrounded radially outwards by a downstream annular portion of an annular platform of the upstream row of mobile vanes. The upstream annular portion of the annular platform of the annular row of stator vanes has a radially external annular face from which fins extend, which are distributed around the longitudinal axis and extend radially outwards towards the downstream annular portion of the platform of the annular row of mobile vanes.

Air-gap fins for a turbine engine compressor

A turbine engine compressor includes a stator featuring an annular casing and at least one annular row of variable-pitch vanes, wherein each vane comprises a radially external end having a pivot mounted in an orifice in the casing and connected by a linking member to an control ring capable of pivoting axially in relation to the casing, where the linking member comprises a first end fixed to the pivot of the vane and a second end having a pin inserted in a hole in the control ring, where at least one of the holes in the control ring serves for insertion of the pins of the linking members, is oblong in shape and extends in the circumferential direction in order to allow movement of the pin in the oblong hole, during rotation of the control ring.

Turbine engine compressor, in particular of an aeroplane turboprop or turbofan

The invention relates to an assembly comprising a turbine engine casing and a bladed rotor wheel arranged therein. The casing presents an inside wall having a circumferential strip of abradable material. Facing the tips of the blades, the casing presents upstream the strip of abradable material, and downstream a circumferential groove. The strip of abradable material is defined downstream by the circumferential groove. The downstream end of the circumferential groove is in register with or downstream from the trailing edges of the blades. This arrangement optimizes the use of abradable material in the casings of the turbine engines.

Growing community of inventors

Moissy-Cramayel, France

Sébastien Claude Cochon