Abstract: A moulding unit for producing a turbomachine blade blank. The moulding unit has an upstream portion and a downstream portion. According to the invention, the moulding unit comprises a raised portion designed to form a downstream opening which, at the level of the trailing edge, passes through at least one of the side edges of the basin-like portion at the blade tip. The raised portion rigidly secures the upstream portion to the downstream portion and comprises a strengthening protuberance extending in the transverse direction.

Abstract: A turbine blade including a root bearing an aerofoil extending in a direction of span and ending in a tip, this aerofoil including a leading edge and a trailing edge which are connected by a pressure face wall and a suction face wall, the trailing edge including on the pressure face side cooling slots. The trailing edge includes one or more first slots close to the root, which are supplied via a lower cavity; one or more last slots near the tip which are supplied via an upper cavity; intermediate slots situated between the first slots and the last slots supplied with air via a downstream line set; and wherein the lower cavities and upper cavities and the downstream line set are supplied distinctly at the level of the root.

Abstract: A method of manufacturing a core for molding a blade of a turbomachine, including placing a flexible ceramic sheet in a cavity of a mold in such a way as to shape the ceramic sheet, introducing a ceramic paste into the cavity of the mold, the ceramic paste forming at least one frame in contact with the ceramic sheet, and co-sintering the ceramic sheet and the frame in such a way as to rigidly join together the ceramic sheet and the frame, the frame maintaining the shape of the ceramic sheet given by the mold during the preceding co-sintering phases.

Abstract: A ceramic core used for fabricating a hollow turbine blade for a turbine engine by using the lost-wax casting technique and shaped to constitute the cavities of the blade as a single element, includes, in order to feed the insides of these cavities jointly with cooling air, core portions that are to form first and second lateral cavities and that are connected to a core portion that is to form at least one central cavity, firstly in the core root via at least two ceramic junctions, and secondly at various heights up the core via a plurality of other ceramic junctions of positioning that defines the thickness of the internal partitions of the blade, while also ensuring additional cooling air for predetermined critical zones of the first and second lateral cavities.

Abstract: A foundry core for manufacturing a blade of a turbomachine including a tip section offset, including a core element for forming various internal cavities, the core element including a leading-edge cavity internal core, central cavity internal cores, and a trailing-edge cavity internal core. The internal core for the central cavity adjacent to the internal core for the trailing-edge cavity includes a bulge extending toward the core for the leading-edge cavity.

Abstract: The internal cooling of moving blades of turbines in aircraft turbomachines is limited in effectiveness because of inhomogeneities of this cooling on each of the pressure-side and suction-side walls. To address this problem, there is proposed a blade including a circuit for cooling the airfoil part thereof, in which circuit the cavities interconnected in series are such that the stream of air flows radially toward the outside along the pressure-side wall in pressure-side cavities, and radially toward the inside along the suction-side wall in a suction-side cavity that is separated from the pressure-side cavities by an internal wall of the airfoil part. In this way, the force of the Coriolis effect deflects the stream of air toward each of the pressure-side and suction-side walls thereby limiting the inhomogeneity.

Abstract: A turbine blade including a root, a vane extending in a spanwise direction, ending at a tip and including a leading edge and a trailing edge and a pressure-side wall and a suction-side wall, the vane further including at least one upstream duct configured to collect air at the root to cool the leading edge, discharging the air through holes passing through the wall of the leading edge; at least one downstream duct separate from the upstream duct and configured to collect air at the root to cool the trailing edge, discharging the air through holes passing through the pressure wall upstream from the trailing edge; an inner side cavity running along the pressure-side wall to form a heat shield insulating the downstream duct.

Abstract: A turbine blade for a turbomachine such as a turboprop or a turbojet, the blade including a root, a vane carried by the root and extending in a spanwise direction, ending at a tip, the vane including a pressure-side wall and a suction-side wall spaced apart from each other, and at least one duct configured to collect cooling air at the root and to make same flow in the vane in order to cool same; holes and/or slots made in the walls of same in order to discharge the cooling air out of the vane; an upper inner cavity located at the tip of the vane in order to cool same; and wherein at least one of the ducts directly supplies the upper cavity with cooling air collected in the root.

Abstract: A turbine blade for a turbomachine, including a vane extending in a spanwise direction between a root and a tip, a first inner side cavity running along a pressure-side wall and a second inner side cavity running along a suction-side wall. The blade further includes at least one inner central duct configured to collect air intended to cool the vane. The central duct extends between the side cavities, being separated from the side cavities to be at least partially thermally isolated from the pressure-side wall and the suction-side wall. The side cavities communicate with each other in a junction region located downstream from the central duct, over a majority of a height of the central duct in the spanwise direction.

Abstract: A method of manufacturing a core for molding a blade of a turbomachine, including placing a flexible ceramic sheet in a cavity of a mold in such a way as to shape the ceramic sheet, introducing a ceramic paste into the cavity of the mold, the ceramic paste forming at least one frame in contact with the ceramic sheet, and co-sintering the ceramic sheet and the frame in such a way as to rigidly join together the ceramic sheet and the frame, the frame maintaining the shape of the ceramic sheet given by the mold during the preceding co-sintering phases.

Abstract: The internal cooling of moving blades of turbines in aircraft turbomachines is limited in effectiveness because of inhomogeneities of this cooling on each of the pressure-side and suction-side walls. To address this problem, there is proposed a blade including a circuit for cooling the airfoil part thereof, in which circuit the cavities interconnected in series are such that the stream of air flows radially toward the outside along the pressure-side wall in pressure-side cavities, and radially toward the inside along the suction-side wall in a suction-side cavity that is separated from the pressure-side cavities by an internal wall of the airfoil part. In this way, the force of the Coriolis effect deflects the stream of air toward each of the pressure-side and suction-side walls thereby limiting the inhomogeneity.

Abstract: A foundry core for manufacturing a blade of a turbomachine including a tip section offset, including a core element for forming various internal cavities, the core element including a leading-edge cavity internal core, central cavity internal cores, and a trailing-edge cavity internal core. The internal core for the central cavity adjacent to the internal core for the trailing-edge cavity includes a bulge extending toward the core for the leading-edge cavity.

Abstract: A turbine blade, including an intrados wall, an extrados wall, at least one first radial trailing edge cavity, at least one second radial cavity upstream of the trailing edge cavity, an internal wall separating the radial cavities and with at least one channel connecting the cavities to one another, the channel being oriented in an axis intersecting the internal surface of the intrados wall.

Abstract: A moving blade for a turbomachine with a central portion that is geometrically subdivided into four adjacent pressure-side zones disposed on the pressure side of the blade, and into four adjacent suction-side zones disposed on the suction side of the blade, the pressure-side and suction-side zones being distributed on either side of the skeleton of the blade, and the blade including in its central portion a pressure-side cooling circuit and a suction-side cooling circuit, the pressure-side cooling circuit including three radial cavities occupying three adjacent pressure-side zones, and the suction-side cooling circuit including three radial cavities occupying the four suction-side zones and the remaining pressure-side zone.

Abstract: A moving blade for a turbomachine, the central portion of the blade including a pressure-side cooling circuit and a suction-side cooling circuit. The pressure-side circuit comprises at least first and second pressure-side cavities extending from the pressure side of the blade to a central wall, a central cavity extending from the pressure side to the suction side of the blade, and outlet orifices opening out from the central cavity and into the pressure-side face of the blade. The suction-side circuit comprises at least first and second suction-side cavities extending radially from the suction side of the blade to the central wall, a central cavity extending across the blade from the pressure side to the suction side, and outlet orifices opening out from the central cavity and into the pressure-side face of the blade.

Abstract: The present invention relates to the field of turbine blades for a turbo machine, in particular a turbine blade (1), including an intrados wall (2), an extrados wall (3), at least one first radial trailing edge cavity (4), at least one second radial cavity (5) upstream of the trailing edge cavity (4), an internal wall (6) separating the radial cavities (4 and 5) and comprising at least one channel (7) connecting the cavities (4 and 5) to one another, the said channel (7) being oriented in an axis (71) intersecting the internal surface (42) of the intrados wall (2).

Abstract: The invention relates to moving blade for a turbomachine. The central portion of the blade is geometrically subdivided into four adjacent pressure-side zones disposed on the pressure side of the blade, and into four adjacent suction-side zones disposed on the suction side of the blade, the pressure-side and suction-side zones being distributed on either side of the skeleton of the blade, and the blade including in its central portion a pressure-side cooling circuit and a suction-side cooling circuit, the pressure-side cooling circuit comprising three radial cavities occupying three adjacent pressure-side zones, and the suction-side cooling circuit comprising three radial cavities occupying the four suction-side zones and the remaining pressure-side zone.

Abstract: A moving blade for a turbomachine, the central portion of the blade including a pressure-side cooling circuit and a suction-side cooling circuit. The pressure-side circuit comprises at least first and second pressure-side cavities extending from the pressure side of the blade to a central wall, a central cavity extending from the pressure side to the suction side of the blade, and outlet orifices opening out from the central cavity and into the pressure-side face of the blade. The suction-side circuit comprises at least first and second suction-side cavities extending radially from the suction side of the blade to the central wall, a central cavity extending across the blade from the pressure side to the suction side, and outlet orifices opening out from the central cavity and into the pressure-side face of the blade.

Abstract: A gas turbine blade comprising an internal cooling circuit consisting of at least one cavity extending radially between the base and tip of the blade, at least one air inlet aperture at one radial end of the cavity and at least one air outlet orifice opening into the cavity and emerging onto one of the faces of the blade. At least one of the walls of the cavity of the cooling circuit comprises at least one air deflector whereof the shape and dimensions are adapted to project the air flowing along the wall of the cavity towards an opposite wall of the cavity whilst avoiding re-attachment of the boundary layer immediately downstream of the air deflector.

Abstract: A gas turbine blade comprising an internal cooling circuit consisting of at least one cavity extending radially between the base and tip of the blade, at least one air inlet aperture at one radial end of the cavity and at least one air outlet orifice opening into the cavity and emerging onto one of the faces of the blade. At least one of the walls of the cavity of the cooling circuit comprises at least one air deflector whereof the shape and dimensions are adapted to project the air flowing along said wall of the cavity towards an opposite wall of said cavity whilst avoiding re-attachment of the boundary layer immediately downstream of said air deflector.