Abstract: The present invention relates to a two-spool gas turbine engine including an HP turbine stator ring and an exterior wall of the transition channel between the HP and LP stages, a first enclosure for controlling the stator ring, and a second enclosure for distributing air for blowing the exterior wall of the transition channel. The engine is characterized in that the two enclosures are placed in communication via an orifice controlled by a valve adapted to be open when the pressure P1 in the first enclosure is greater than the pressure P2 in the second enclosure, and closed when P1<P2.

Abstract: An inter-turbine casing for a turbofan is disclosed. The casing includes an outer ring, an inner ring, and an intermediate ring between the inner ring and the outer ring. The inner and intermediate rings have respective openings for the passage of cooling air. The casing includes at least one sealing device, arranged between the outer ring and the intermediate ring, with a base, provided with at least one orifice for the passage of cooling air, and a peripheral skirt which is able to be compressed and expanded elastically. The base bears against the outer ring and the peripheral skirt bears against the intermediate ring.

Abstract: A cooled blade forming an upstream guide vane element for a turbomachine is disclosed. The airfoil includes a longitudinal cavity with a first opening at one end and a second opening at the other end, a tubular sleeve being housed in the cavity with a first end in the first opening and a second end in the second opening, first spacers on the side of the first end, and second spacers on the side of the second end of the sleeve making a space between the outer face of the sleeve and the wall of the cavity. The blade is arranged so that the sleeve is inserted into the cavity through the first opening. The first spacers are secured to the sleeve and the second spacers are secured to the wall of the cavity of the airfoil. The invention makes it possible to mount the sleeve despite an accentuated curvature of the profile of the airfoil.

Abstract: The monocrystalline turbomachine blade according to the invention that is cast and directionally solidified, is disclosed. The blade includes an airfoil with a leading edge, a pressure face, a suction face, a trailing edge, a skeleton and having a longitudinal axis, the faces and having a neck line, respectively a pressure face neck and a suction face neck relative to the adjacent blade in the turbomachine rotor of which it forms an element; an endpiece of the airfoil, such as a heel or a platform, having an airfoil end face, on the stream side, forming an angle with the axis ZZ; and a connection zone between the airfoil and the airfoil end face. The connection zone forms a fattening of the airfoil. The connection zone extends about the leading edge between a point P1 situated on the suction face of the airfoil upstream of the suction face neck and a point P3 situated on the pressure face of the airfoil upstream of the pressure face neck.

Abstract: A transition channel between a first turbine section and a second turbine section for a gas turbine engine is disclosed. The channel includes a first radially external annular wall including ring sector elements housed inside an annular ring and orifices for injecting a fluid into the channel in order to re-energize the boundary layer thereof, a second radially internal annular wall, and a fluid supply device arranged between the space outside the ring and the injection orifices.

Abstract: A device for cooling a turbine casing in a turbomachine including a turbine is disclosed. The turbine includes several stages, at least one of the stages includes a nozzle assembly formed of an annular row of fixed vanes and an impeller mounted to rotate in a cylindrical shroud formed of ring sectors fixed to the casing, a cooling circuit including ducts carrying cooling air into cavities formed in the vanes of the nozzle assembly, and an air-carrying arrangement which carries air to casing upstream hooks for suspending the ring sectors.

Abstract: A joining device including a clevis and a connecting member is disclosed. The connecting member includes a through hole which is housed between a first and a second lug of the clevis. These lugs respectively have a first clevis hole and a second clevis hole. The device further includes a shear pin system passing through each of the holes and includes a threaded end collaborating with a nut system which abuts against the second lug. The shear pin system includes a widened portion passing through the first hole and forming a shoulder situated between the two lugs. The member is situated away from the first lug. A first side of the member is pressed against the shoulder and a second side of the member is pressed against the second lug.

Abstract: The present invention relates to a two-spool gas turbine engine including an HP turbine stator ring and an exterior wall of the transition channel between the HP and LP stages, a first enclosure for controlling the stator ring, and a second enclosure for distributing air for blowing the exterior wall of the transition channel. The engine is characterized in that the two enclosures are placed in communication via an orifice controlled by a valve adapted to be open when the pressure P1 in the first enclosure is greater than the pressure P2 in the second enclosure, and closed when P1<P2.

Abstract: The monocrystalline turbomachine blade according to the invention that is cast and directionally solidified, comprising: an airfoil with a leading edge, a pressure face, a suction face, a trailing edge, a skeleton and having a longitudinal axis, the faces and having a neck line, respectively a pressure face neck and a suction face neck relative to the adjacent blade in the turbomachine rotor of which it forms an element, an endpiece of the airfoil, such as a heel or a platform, having an airfoil end face, on the stream side, forming an angle with the axis ZZ and a connection zone between the airfoil and said airfoil end face, said connection zone forming a fattening of the airfoil, is characterized in that said connection zone extends about the leading edge between a point P1 situated on the suction face of the airfoil upstream of the suction face neck and a point P3 situated on the pressure face of the airfoil upstream of the pressure face neck.

Abstract: The present invention relates to a cooled blade forming an upstream guide vane element for a turbomachine, wherein the airfoil comprises a longitudinal cavity with a first opening at one end and a second opening at the other end, a tubular sleeve being housed in the cavity with a first end in the first opening and a second end in the second opening, first spacers on the side of the first end and second spacers on the side of the second end of the sleeve making a space between the outer face of the sleeve and the wall of the cavity, the blade being arranged so that the sleeve is inserted into the cavity through the first opening. The blade is characterized in that the first spacers are secured to the sleeve and the second spacers are secured to the wall of the cavity of the airfoil. The invention makes it possible to mount the sleeve despite an accentuated curvature of the profile of the airfoil.

Abstract: The present invention relates to a joining device (10) comprising a clevis (22) and a connecting member (28), the member (28) pierced with a through hole (30) being housed between a first and a second lug (26a, 26b) of the clevis, these lugs respectively having a first clevis hole (32a) and a second clevis hole (32b), the device further comprising a shear pin system (36) passing through each of said holes (32a, 32b, 30) and comprising a threaded end (46) collaborating with a nut system (44) bearing against the second lug. According to the invention, the shear pin system comprises a widened portion (42) passing through the first hole and forming a shoulder (50) situated between the two lugs, the member situated some distance from the first lug being pressed against the shoulder (50) on the one hand and against the second lug (26b) on the other.

Abstract: The invention relates to an inter-turbine casing (40) for a turbofan (2), comprising an outer ring (44), an inner ring (54), and an intermediate ring (52) between the inner ring and the outer ring, the inner (54) and intermediate (52) rings having respective openings (540, 520) for the passage of cooling air. The casing is distinguished in that it comprises at least one sealing device (80), arranged between the outer ring (44) and the intermediate ring (52), with a base (82), provided with at least one orifice (90) for the passage of cooling air, and a peripheral skirt (84) which is able to be compressed and expanded elastically, the base (82) bearing against the outer ring (44) and the peripheral skirt (84) bearing against the intermediate ring (52).

Abstract: The present invention relates to a transition channel (10) between a first turbine section and a second turbine section for a gas turbine engine, comprising a first radially external annular wall (102), a second radially internal annular wall, the first wall comprising orifices (102C?) for injecting a fluid into the channel in order to re-energize the boundary layer thereof. It is characterized in that the first wall (102) consists of ring sector elements housed inside an annular ring (104), fluid supply means being arranged between the space (110) outside the ring and said injection orifices (104).

Abstract: A connection device between a nozzle and a feed enclosure for feeding cooling fluid to at least one injector in a turbomachine, the device comprises at least one bushing having open axial ends engaged in orifices in a wall of the nozzle and in a wall of the feed enclosure, together with retaining means for retaining the bushing axially, the retaining means being formed to project from an outer cylindrical surface of the bushing between the axial ends of the bushing and at a distance from said ends.

Abstract: Device for cooling a turbine casing in a turbomachine, this turbine comprising several stages each including a nozzle assembly (18) formed of an annular row of fixed vanes (20) and an impeller mounted to rotate in a cylindrical shroud formed of ring sectors (34) fixed to the casing, and a cooling circuit comprising ducts carrying cooling air into cavities (46) formed in the vanes (20) of the nozzle assembly, and means for carrying air to casing upstream hooks (72) for suspending the ring sectors (34).

Abstract: A strut (1) fixed to a main portion of turbojet stator carries a ring (5) by a pair of hooks (4, 6) on one side and joining lips (7 and 8) on the other side. In accordance with the invention, the seal is made on the flat sides (10, 11) adjacent to the lips by elastic deflection of the tabs (13) pushing back one of the lips (7). The tabs (13) can also determine the tangential positioning of the ring (5).

Abstract: Distributor sectors (25) on which fixed vanes are placed adjacent to a casing (31) by means of sealing sectors (32) alternating with them in the axial direction and provided with force resistant faces (37; 39) through which forces exerted on the distributor sectors are transmitted to the casing. The inner surface of the casing (31) is smoother and is not fitted with any hooks; consequently, this casing (31) is less complicated to make and is less mechanically stressed to retain the distributor sectors, and the distributor sectors may be installed by a purely axial movement.

Abstract: A strut (1) fixed to a main portion of turbojet stator carries a ring (5) by a pair of hooks (4, 6) on one side and joining lips (7 and 8) on the other side. In accordance with the invention, the seal is made on the flat sides (10, 11) adjacent to the lips by elastic deflection of the tabs (13) pushing back one of the lips (7). The tabs (13) can also determine the tangential positioning of the ring (5).

Abstract: A thrust reverser is disclosed for a high bypass ratio turbofan jet engine having a fan cowling forming an outer boundary of an airflow duct and a jet engine structure including a jet engine cowling. The thrust reverser has a casing affixed to the jet engine structure, the casing having a plurality of attaching yokes. A first thrust reverser flap is pivotally connected to a corresponding one of the attaching yokes and is located rearwardly of an exit of the airflow duct, each first thrust reverser flap having a distal end and being movable between a forward thrust position, and a reverse thrust position. A second thrust reverser flap is pivotally connected to the distal end of a corresponding one of the first thrust reverser flaps and is also movable between a forward thrust position, and a reverse thrust position, wherein the second thrust reverser flaps also redirect air flow emanating from the air flow duct.