Societe Nationale D'etude Et De Construction De Moteurs D'aviation

Carmen Miraucourt

Remy P Ritt

Gilles L Delrieu

Georges Mazeaud

Pascal Noel Brossier

Jean-Francois R Imbault

Didier Marie Mortgat

Christian Kaincz

Roland R Mestre

Philippe Thieux

A device for suspending a turboprop engine from a structural element of an aircraft comprises a suspension pylon parallel to the longitudinal axis of the engine and fixed to the structural element, resilient connections between the engine and the pylon and an arch fixed to the pylon, and a torsion bar carried by the arch so that the bar can pivot freely about its own axis, the opposite ends of the torsion bar being connected to opposite sides of the engine by respective lever and link rod assemblies so that the torsion bar takes up the reaction torque about the longitudinal axis of the engine.

Suspension device for a turboprop engine

A mounting system for a stator vane assembly is disclosed for mounting a vane and its cooling system to minimize the thermal and mechanical stresses imparted to the vane structure. Platforms are attached to opposite ends of the vane, which has a hollow interior and a cooling gas distributor located within the interior of the vane. One end of the cooling gas distributor is also fixedly attached to one of the vane platforms, while the opposite end of the cooling gas distributor is attached to a casing, not to the opposite vane platform. Spaces are defined between the cooling gas distributor and the vane platform, as well as the casing and the vane platform. A gas inlet sleeve may be attached to the cooling distributor and serve as its attachment to the external casing. The gas inlet sleeve extends through an opening in the vane platform which may have a greater diameter than that of the sleeve in order to provide a cooling passage interconnecting the spaces between the vane platform and the cooling gas distributor, and the vane platform and the casing.

Mounting system for a stator vane

A hollow turbine blade equipped with a cooling system comprises a pair of internal partitions connected to the intrados and extrados walls of the blade and dividing the interior of the blade into at least one internal cavity extending from the root end to the tip end of the blade, a median partition connected to the two internal partitions to divide the cavity defined between the two internal partitions into two half-cavities, the median partition leaving an aperture at or near one end of the blade to establish communication between the two half-cavities, air inlet holes opening into the half-cavity on the intrados side of the median partition at or near the other end of the blade, air flow disturbing regions disposed on at least part of the internal face of the intrados wall opposite the median partition, and air outlet holes provided in the extrados wall opposite the median partition. Preferably each half-cavity is divided into two passages by half-partitions extending between the median partition and the intrados and extrados walls from root to tip of the blade.

Hollow turbine blade with internal cooling system

The present invention relates to a compressor casing for a gas turbine engine in which an inner casing is radially adjustable so as to maintain a radial clearance between the casing and a rotor as the rotor undergoes radial expansion and contraction. The compressor casing has an outer casing and a concentrically arranged inner casing with a plurality of generally radially extending arms interconnecting the inner and outer casings. The volume between the casings is sealed so as to form a chamber and heated air bled from a downstream stage of the compressor is directed into this chamber such that the heated air contacts the exterior surfaces of the hollow radial arms. The invention also has a system for supplying cooling air to the interior of each of the radial hollow arms including a regulating device so that the cooling air can be selectively applied to the arms.

Compressor casing for a gas turbine engine

A compressor casing in a gas turbine engine includes an inner jacket formed by two half-shells connected by coupling flanges having cylindrical bellows fitted therebetween and arranged to expand and contract to control the inner diameter of the casing depending on the pressure within the bellows as provided by a supply of air taken from the downstream end of the compressor. The two half-shells are also interconnected by an additional flexible connection.

Gas turbine engine compressor casing with internal diameter control

A turbine ring comprises an annular support in two parts and a ring of ceramic sectors for sealing purposes. The two parts of the support are interconnected by an axial groove and a sliding male part. Each part comprises an axial annular groove in which is engaged a respective axial edge portion of each sector. The cooperating internal radial faces respectively of the edge portion and of the groove are circumferential, while the radially outer radial face of each sector comprises at least one flat zone. The corresponding face of each groove may be polygonal.

Turbine ring incorporating elements of a ceramic composition divided

A turbine ring assembly is mounted on an annular carrier secured to the irior of the casing of a turbine and includes a succession of contiguous sectors of a ceramic abradable material. The sectors have grooves in which support fingers are engaged rigid with an annular member connected to the casing of the turbine. The sectors are thereby held in an annular configuration and the support fingers are so located in the grooves as to provide a screen to prevent reintroduction of hot gases.

Growing community of inventors

Brie Comte Robert, France

Carmen Miraucourt