A nozzle module includes a nozzle body having a blade shape in a cross section and extending in a radial direction, and a platform member integrally connected to each end portion of the nozzle body in the radial direction. The platform member includes a first portion formed on a first side in an axial direction in which a central axis extends, and having a pair of first side surfaces extending in the axial direction, when viewed in the radial direction, and a second portion formed to extend to a second side in the axial direction with respect to the first portion, and having a second side surface extending obliquely with respect to the first side surface, when viewed in the radial direction.

A turbomachine defines an axial direction, a radial direction perpendicular to the axial direction, and a circumferential direction extending concentrically around the axial direction. The turbomachine includes a nozzle having an inner platform, an outer platform, and an airfoil. The airfoil includes a leading edge, a trailing edge downstream of the leading edge, a pressure side surface, and a suction side surface opposite the pressure side surface. The trailing edge defines a circular arc between the inner platform and the outer platform.

A turbomachine defines an axial direction, a radial direction perpendicular to the axial direction, and a circumferential direction extending concentrically around the axial direction. The turbomachine includes a nozzle having an inner platform, an outer platform, and an airfoil. The airfoil includes a leading edge, a trailing edge downstream of the leading edge, a pressure side surface, and a suction side surface opposite the pressure side surface. The trailing edge defines a circular arc between the inner platform and the outer platform.

This blade repair method has: a first welding step in which overlay welding in which a first welding material is used is performed to form a notched part and a bury a first region positioned on a blade-body side with a first welding material; and a second welding step in which, after the first welding step, overlay welding in which a second welding material is used is performed to form a notched part and bury a second region positioned on a front-surface side of a platform with the second welding material. The high-temperature strength of the second welding material is higher than the high-temperature strength of the first welding material, the weldability of the first welding material is higher than the weldability of the second welding material, and the second region is located in a range from 1.0 mm to 3.0 mm (inclusive) from the front surface of the platform toward the blade body.

A turbine is provided with: a turbine wheel configured to rotate about an axis O1; a turbine housing accommodating the turbine wheel and defining an annular nozzle passage on the outer peripheral side of the turbine wheel; and a plurality of low solidity nozzle vanes 6 arranged in the nozzle passage at an interval in the circumferential direction. Circumferentially adjacent low solidity nozzle vanes 6 are disposed at different radial positions in a connection position of each of the low solidity nozzle vanes 6 with a hub-side wall surface of the hub-side wall surface and a shroud-side wall surface which define the nozzle passage.

A turbine module (2) for a turbomachine (1). The turbine module (2) includes a main channel (26) to guide a main flow (36) through the turbine module (2), a rotor blade (21) and a stator vane (22), the stator vane (22) including a stator airfoil (22) and a platform (23), with the stator airfoil (22) arranged downstream of the rotor blade (21) in the main channel (26), and a cavity (30) including an inlet (31) for injecting a part (36.2) of the main flow (36) into the cavity (30), an outlet (32) for a reinjection of the part (36.2) of the main flow (36) from the cavity (30) into the main channel (26), wherein the cavity (30) is arranged at an axial position of the stator vane (20) and is radially offset from the stator airfoil (22).

An annular intermediate casing for a turbomachine through which an aerodynamic airstream circulates, including a radially internal shroud (31), a radially external shroud, at least one stator blade including a vane with a leading edge, and at least one aero-dynamic member which is formed of a structural arm extending radially at least in part between the radially internal shroud and the radially external shroud. The aerodynamic member includes a leading edge which is aligned with the leading edge of the stator blade in one plane and which is borne by a profiled portion positioned upstream of the structural arm in the direction of circulation of the aerodynamic airstream, the profiled portion, the structural arm, the radially internal shroud and the radially external shroud being monobloc.

An internal core profile for a turbine nozzle airfoil of a gas turbine is provided. The turbine nozzle may include an airfoil core having an uncoated nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, wherein the X, Y, and Z coordinates are distances in inches measured in a

Cartesian coordinate system, the corresponding X and Y coordinates, when connected by a smooth continuous arc, define one of a plurality of airfoil core profile sections at each Z distance, and the plurality of airfoil core profile sections, when joined together by smooth continuous arcs, define an airfoil core shape.

A system for retaining stators and reducing air leakage in a gas turbine engine having an axis includes a stator having an inner platform, an outer platform, a low pressure side, a high pressure side, and at least one foot, and designed to turn air. The system also includes a case positioned radially outward from the stator and having at least one recess designed to interface with the at least one foot to resist movement of the stator relative to the case. The system also includes a bladder positioned between the outer platform of the stator and the case and designed to receive pressurized fluid having a greater pressure than ambient pressures experienced at the low pressure side of the stator and to further resist movement of the stator relative to the case in response to receiving the pressurized fluid.

A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection is between the fan case and the core engine includes three triangular-frame connecting members rigidly connected to the fan case at a fan case connection point, and to the core engine at a core engine connection point. The triangular-frame connecting members each are defined by two rigid legs which extend between the fan case and to the core engine, along directions each have a component extending radially inwardly and a component in opposed circumferential directions to each other. A plurality of non-structural fan exit guide vanes and the non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise. The non-structural fan exit guide vanes are rigidly mounted to at least one of the fan case and the core engine.