Abstract: A guide vane for a dual flow aircraft turbine engine includes an aerodynamic part. The aerodynamic part includes an inner duct for lubricant cooling extending in a main direction and being partly bounded by a pressure side wall and a suction side wall of the vane. The aerodynamic part is embodied as a single piece, and includes heat transfer fins arranged in the duct expanding substantially parallel to a direction of the duct. The fins are spaced from each other depending on the direction of the duct and a transversal direction of the vane, so that at least some of the fins are arranged substantially staggered.

Abstract: A ring of vanes for an aircraft turbine engine, the ring presenting an axial direction and a radial direction and including a first annular sheath presenting an inside surface and a second annular sheath presenting an inside surface facing the first inside surface of the first annular sheath, the first and second sheaths being coaxial and defining between them a flow passage for a gas stream. The ring further includes both partially annular acoustic treatment panels including resonant cavities and also vanes extending in the radial direction between the first and second sheaths. Each vane includes a radial airfoil with at least two attachment tabs at each radial end of the airfoil fastened to the first or second sheath, and the inside surface of at least one of the first and second annular sheaths and the corresponding attachment tabs are covered by acoustic treatment panels arranged between the airfoils.

Abstract: The invention relates to a guide vane for a twin-spool aircraft turbomachine, of which the aerodynamic part (32) comprises an internal lubricant cooling passage (50a) extending along a principal lubricant flow direction (52a). According to the invention, the aerodynamic part is made in a single piece and also comprises heat transfer fins (80a, 80b) arranged in the passage (50a) connecting the intrados and extrados walls (70, 72) and extending approximately parallel to the direction (52a), these fins being distributed in successive rows along the principal direction (52a) and made such that for two rows (R1, R2) of staggered directly consecutive fins, the row (R1) comprises fins (80a) forming a positive acute angle A1 with a dummy reference plane (Pf), while the row (R2) comprises fins (80b) forming a negative acute angle A2 with this plane (Pf).

Abstract: The invention relates to a guide vane intended to be mounted in a turbine engine between an inner shroud (17) and an outer shroud (16), comprising a longitudinal straightening body (41) for an air flow extending between a first end (42) intended to be positioned at the inner shroud (17) and a second end (44) intended to be positioned at the outer shroud (16), the longitudinal straightening body having an aerodynamic profile defined by a leading edge (41a) and a trailing edge (41b) in the flow direction of the air flow, and by a camber line (41c) extending from the leading edge (41a) to the trailing edge (41b).

Abstract: A support structure suitable for interposing between the engine and the nacelle of an aeroengine and suitable for being fastened on an intermediate casing is provided. The structure includes a shroud having a plurality of sectors, a plurality of radial arms each installed between two adjacent shroud sectors, and a plurality of fasteners for fastening the plurality of shroud sectors to the plurality of radial arms. The shroud sectors and radial arms form a skeleton for transmitting structural forces within the intermediate casing. The structure includes non-structural fairings mounted on the skeleton and suitable for reconstituting the airflow passage. Each of the shroud sectors and radial arms is formed by a metal frame without mechanical discontinuity, each shroud sector metal frame defining at least one opening that is closed by a cover and each radial arm metal frame defining at least one orifice for receiving a closure plate.

Abstract: The invention relates to an intermediate case (25) for a twin spool turbomachine for an aircraft, comprising a hub (26), an outer shell (23) and outlet guide vanes (24) installed at their ends on the hub and on the outer shell, and each of at least some of the outlet guide vanes (24) performing a heat exchanger function and comprising a lubricant passage (50a, 50b) designed to be cooled by the fan flow (58) following an outer surface of the outlet guide vane. According to the invention, the case also comprises at least one lubricant duct (55) passing along a circumferential direction of the hub (26) and at least part of which is made from a single casting with the hub, the lubricant duct (55) having at least one lateral opening communicating with the lubricant passage (50a, 50b) of at least one of the vanes (24).

Abstract: A closer for an attachment orifice of an OGV in a turbine engine, the vane having a blade fixed to a root provided with attachment orifices, the closer including a cup adapted to be fixed in the attachment orifice of an OGV; a cylindrical shaped plug engaging in the cup, the plug including an upper face inclined relative to the ground plane of the plug; a guide adapted to position the plug on the cup by rotation into a working position, in which the upper face of the plug coincides with the surface level of the vane root; a blocking system to block the plug in the working position.

Abstract: A guide vane for a dual flow aircraft turbine engine, the aerodynamic part of the vane including an inner duct for lubricant cooling extending in a main direction and being partly bounded by a pressure side wall and a suction side wall of the vane. A heat conduction matrix is lodged in the duct, and presents main heat transfer wings extending parallel to the direction, and laid out in staggered rows.

Abstract: A vaned wheel device traversed by a radial drive shaft includes radial arms, one of which contains the shaft, and stationary vanes manufactured separately from the casing and a ferrule and assembled thereto. A sleeve encompassing the shaft traverses the casing and serves to support a bearing for supporting the shaft with positional accuracy, and maintaining tightness by forming a single cavity with the radial arm cavities traversing the shaft.

Abstract: The invention relates to a guide vane (24) for a dual flow aircraft turbine engine, the aerodynamic part (32) of the vane comprising an inner duct (50a) for lubricant cooling extending in a main direction (52a) and being partly bounded by a pressure side wall (70) and a suction side wall (72) of the vane. According to the invention, the aerodynamic part (32) of the vane is embodied from as a single piece, including also the heat transfer fins (80), arranged in the duct (50a) expanding substantially parallel to the direction (52a), the fins (80) being spaced from each other depending on the direction (52a) as well as a transversal direction (60) of the vane, so that at least some of the fins (80) are arranged substantially staggered.

Abstract: A turbomachine fan frame including a central hub, a plurality of radial vanes installed on the hub and distributed around the hub, each vane including an inner radial end root at which the vane is fixed to the hub, a plurality of flow stream reconstitution platforms, each platform being installed between the roots of two adjacent vanes and radially at a distance from the hub, wherein each platform is fixed to the hub by at least one support element associated with the platform, the support element located between the associated platform and the hub and located between the roots of two adjacent vanes.

Abstract: A plate closes the main part of a vane in two parts. An internal cavity is closed by the plate. Around a bearing surface of the plate, a groove receives an elastic seal, and the outer edge of the plate extends over the groove to compress the seal and to establish excellent leak proofing. The position of the plate may be guaranteed by centring pins penetrating into holes adjacent to the seal and which it fills when the pins have been removed, which completes the closing of the cavity. The plate has application in particular to stationary gas guide vanes in turbine engines.

Abstract: This invention relates to an assembly (100) for an aircraft turbofan comprising a flowpath splitter (40) downstream from which there is firstly a fan flowpath (31) inner delimitation ring (44) and secondly a compressor case (42), the assembly comprising a plurality of outlet guide vanes (34) provided with an inner platform (50) located in the downstream continuity of the delimitation ring. According to the invention, the assembly comprises at least one part (52) to support the delimitation ring (44), comprising a radially outer end (52b) fixed to a downstream part of this ring (44) and a radially inner end (52a) fixed to the compressor case (42).

Abstract: A support structure suitable for interposing between the engine and the nacelle of an aeroengine and suitable for being fastened on an intermediate casing, the structure comprising a shroud having a plurality of sectors (12A, 12B, 12C, 12D), a plurality of radial arms (14A, 14B, 14C, 14D) each installed between two adjacent shroud sectors, and a plurality of fastener means (16) for fastening the plurality of shroud sectors to the plurality of radial arms, the shroud sectors and radial arms forming a skeleton (10) suitable for transmitting structural forces within the intermediate casing, the structure including non-structural fairings mounted on the skeleton and suitable for reconstituting the airflow passage.

Abstract: A closer for an attachment orifice of an OGV in a turbine engine, the vane having a blade fixed to a root provided with attachment orifices, the closer including a cup adapted to be fixed in the attachment orifice of an OGV; a cylindrical shaped plug engaging in the cup, the plug including an upper face inclined relative to the ground plane of the plug; a guide adapted to position the plug on the cup by rotation into a working position, in which the upper face of the plug coincides with the surface level of the vane root; a blocking system to block the plug in the working position.

Abstract: This invention relates to an assembly (100) for an aircraft turbofan comprising a flowpath splitter (40) downstream from which there is firstly a fan flowpath (31) inner delimitation ring (44) and secondly a compressor case (42), the assembly comprising a plurality of outlet guide vanes (34) provided with an inner platform (50) located in the downstream continuity of the delimitation ring. According to the invention, the assembly comprises at least one part (52) to support the delimitation ring (44), comprising a radially outer end (52b) fixed to a downstream part of this ring (44) and a radially inner end (52a) fixed to the compressor case (42).

Abstract: A plate closes the main part of a vane in two parts. An internal cavity is closed by the plate. Around a bearing surface of the plate, a groove receives an elastic seal, and the outer edge of the plate extends over the groove to compress the seal and to establish excellent leak proofing. The position of the plate may be guaranteed by centring pins penetrating into holes adjacent to the seal and which it fills when the pins have been removed, which completes the closing of the cavity. The plate has application in particular to stationary gas guide vanes in turbine engines.

Abstract: A structural and aerodynamic module including two blades each extending along a longitudinal axis, in which each blade includes a leading edge, a trailing edge, a radially inner edge, and a radially outer edge, is provided. The module also includes a radially inner platform mechanically connecting the radially inner edges of the two blades, and a radially outer platform mechanically connecting the radially outer edges of the two blades such that the two blades and the radially outer and inner platforms delimit a flow stream. The two blades and the platforms are made from a composite material and the two blades are fixed on the radially outer and inner platforms using screw-nut assemblies.

Abstract: A vaned wheel device traversed by a radial drive shaft includes radial arms, one of which contains the shaft, and stationary vanes manufactured separately from the casing and a ferrule and assembled thereto. A sleeve encompassing the shaft traverses the casing and serves to support a bearing for supporting the shaft with positional accuracy, and maintaining tightness by forming a single cavity with the radial arm cavities traversing the shaft.

Abstract: A structural frame for a turbomachine, such as an intermediate frame or an exhaust frame, the frame including at least two coaxial rings, an inner ring and an outer ring, together with radial arms interconnecting the rings. At least one of these elements includes a plurality of parts that are fastened together by bolting.