Various embodiments of the invention include turbine nozzles and systems employing such nozzles. Various particular embodiments include a turbine nozzle having: an airfoil having: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side; and at least one endwall connected with the airfoil along the suction side, pressure side, trailing edge and the leading edge, the at least one endwall including a non-axisymmetric contour proximate a junction between the endwall and the leading edge of the airfoil.

Various embodiments of the invention include turbine nozzles and systems employing such nozzles. Various particular embodiments include a turbine nozzle having: an airfoil having: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side; and at least one endwall connected with the airfoil along the suction side, pressure side, trailing edge and the leading edge, the at least one endwall including a non-axisymmetric contour proximate a junction between the endwall and the leading edge of the airfoil.

A blade cascade of a turbomachine having at least one shape variation of a blade situated on the blade side in the proximity of a side wall and extending downstream, and at least one side wall contouring of the side wall or at least one second shape variation of an adjacent blade near the side wall, as well as a turbomachine, are disclosed.

The present invention thus proposes an inlet cone for an aircraft turbine engine, comprising a frustoconical body and a tip made from elastically deformable material fixed to an end of smaller diameter of said body, the tip comprising a top configured to be situated on an axis of rotation of the cone and a fastening base for attachment on said end of said body. Said base extends in a connecting plane P. Said connecting plane P is inclined relative to said axis of rotation. Said base has a generally circular or oval shape. According to the invention, said connecting plane P is inclined relative to a transverse plane T perpendicular to said axis of rotation.

A method for producing a rotor disk or a rotor blade for a gas turbine compressor stage or turbine stage of an aeroengine, wherein at least one blade groove of the rotor disk for arrangement of a blade foot of a rotor blade for fastening the rotor blade to the rotor disk, or a blade foot of the rotor blade for arrangement in a blade groove of a rotor disk for fastening the rotor blade to the rotor disk is fabricated using an electrochemical material removal and in the axial direction has a profile which is curved once or more.

The invention relates to a de-icing device for an air intake of an aircraft turbojet engine nacelle extending along an X-axis in which an air stream flows from upstream to downstream, the air intake having an inner cavity, extending annularly about the X-axis, which comprises an inner wall facing the X-axis and an outer wall which is opposite the inner wall, the walls being connected by a leading edge, the de-icing device comprising at least one injector for injecting a stream of hot air into the inner cavity, the injector comprising a nozzle extending along a nozzle axis, the nozzle being configured to inject a stream of hot air having a dissymmetry along the nozzle axis.

An assembly forming an acoustic insulator having a first sheet, a pierced second sheet, and a plurality of first and second structures. Each first structure comprises a first and a second strip, wherein each is shaped to form half of the wall of a cage and wherein, for two successive halves, each strip comprises a facet of a joining wall. Each second structure is made up of a first and a second strip, wherein each is shaped to form half of the wall of a cone, wherein. For each strip, at least one of the wall halves of each cone is pierced. For two successive halves, each strip comprises one facet of the connecting wall. Each cone is located in a cell and each connecting wall is located between the two facets of a joining wall, and, between two adjacent first structures, a second structure is likewise fitted.

A gas turbine engine includes, among other things, a bearing compartment, a shaft configured to rotate during operation of the gas turbine engine, and a sleeve configured to rotate with the shaft during operation of the gas turbine engine. The sleeve includes a variable radial dimension about a circumference of the sleeve. The engine further includes a seal statically mounted relative to the sleeve and configured to cooperate with the sleeve to seal the bearing compartment.

The present invention thus proposes an inlet cone for an aircraft turbine engine, comprising a frustoconical body and a tip made from elastically deformable material fixed to an end of smaller diameter of said body, the tip comprising a top configured to be situated on an axis of rotation of the cone and a fastening base for attachment on said end of said body. Said base extends in a connecting plane P. Said connecting plane P is inclined relative to said axis of rotation. Said base has a generally circular or oval shape. According to the invention, said connecting plane P is inclined relative to a transverse plane T perpendicular to said axis of rotation.

A turbine bucket may include a platform, an airfoil extending from the platform at an intersection thereof, and a cooling circuit extending within the platform and the airfoil. The cooling circuit may include a root turn with an asymmetric shape to reduce stress concentrations therein. The asymmetric shape of the root turn may be asymmetrical along a path between a pressure side of the airfoil and a suction side of the airfoil. The asymmetric shape of the root turn may be asymmetrical within a plane defined by a radial direction and a circumferential direction.