A rotor stage (10) of a turbine engine includes a circumferential row of rotor segments (12), each including: first and second endwalls (14, 16) spaced apart radially, and a first and second sidewalls (18, 20) extending radially between the first and second endwalls (14, 16) and spaced apart circumferentially. The first and second endwalls (14, 16) and the first and second sidewalls (18, 20) define therewithin a flow passage (22) for hot gas. Circumferentially adjacent segments (12a, 12b) mate along a respective split-line (24) extending along an interface between the first sidewall (18) of a first segment (12a) and the second sidewall (20) of a second circumferentially adjacent segment (12b). A composite airfoil structure (26) is thereby defined having a pressure sidewall (18) formed by the first sidewall (18) of the segment (12a) and a suction sidewall (20) formed by the second sidewall (20) of the second segment (12b). The first and second endwalls (14, 16) are respectively configured as a platform (14) and a tip shroud (16) of the segment (12).

The invention relates to a blade assembly 30 for a large dimension axial fan 32 having a rotation axis X. The blade assembly of the invention comprises: a root structure 34 intended to mechanically connect the blade assembly to a hub 36; a blade, wherein at least one portion of the blade has a composite airfoil 46 comprising a fore semi-airfoil 48 and an aft flap 50, wherein: the semi-airfoil is intended to be assembled at a predefined pitch angle α.sub.c with respect to the hub 36 by means of the root structure; the flap 50 is mounted on the blade such that it can be fixed in a position comprised between a maximum deflection position and a minimum deflection position with respect to the pitch angle α.sub.c; and between the fore semi-airfoil and the aft flap a channel 54 is defined suitable for allowing a fluid flow from the face v to the back d of the composite airfoil. The invention further relates to a fan comprising a plurality of blades.

A turbine blade having a gas film cooling structure with a composite irregular groove. The turbine blade has a hollow structure, and a plurality of first grooves which are recessed grooves are provided on an outer surface thereof. A plurality of discrete holes A extending to an inner surface of the turbine blade are provided at the groove bottom of each first groove. The first groove is an irregular groove, and includes at least two portions in a depth direction. A portion having a depth H.sub.1 from the groove bottom of the first groove is a first portion, and the rest thereof is a second portion. At least one side surface of the second portion is formed by expanding in lateral direction from a corresponding side surface of the first portion.

A method of manufacturing a compressor stator having: a first stator blade with a first leading edge and a first trailing edge; a second stator blade disposed a circumferential distance from the first stator blade, the second stator blade having a second leading edge disposed an axial distance from the first leading edge and a second trailing edge disposed an axial distance from the first trailing edge; the method comprising: using additive manufacturing to deposit and fuse together progressive layers of metal material commencing at a substrate to form the first stator blade, the second stator blade, at least one intermediate support structure disposed between the first stator blade and the second stator blade, and at least one primary support structure disposed between the substrate and at least one of: the first stator blade; and the second stator blade; and removing the primary support structure and the intermediate support structure.

Gas turbine aircraft engine comprising an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor; and a fan upstream of the engine core and driven by the turbine, the fan comprising a circumferential row of tandem fan blades. Each fan blade comprises a main blade and an auxiliary blade. Over substantially all of the auxiliary blade's radial span, the leading edge of the auxiliary blade is rearwards of the closest point on the trailing edge of the main fan blade, and on a given aerofoil chordal section of the main fan blade, the leading edge position of an aerofoil chordal section of the auxiliary fan blade lies on a rearwards extension of the camber line of the aerofoil chordal section of the main fan blade, and the main fan blade and the auxiliary fan blade are arranged to rotate in tandem.

A turbine rear structure for a gas turbine engine includes a central hub and a circumferential outer ring coaxial with the central hub. The turbine rear structure further includes a plurality of guide vanes extending radially between the central hub and the circumferential outer ring, and an intermediate guide vane located in a space defined between adjacent guide vanes. The intermediate guide vane is located closer to one of the guide vanes than the other guide vane.

An impeller rotatable about a rotation axis, has a hub, blades extending from the hub to tips circumferentially distributed around the rotation axis, the blades having pressure sides and suction sides, the blades including splitter blades interspersed between full blades, a chord length of the splitter blades less than a chord length of the full blades, first flow channels defined between pressure sides of the splitter blades and suction sides of the full blades, second flow channels defined between suction sides of the splitter blades and pressure sides of the full blades, respective widths of the first and second flow channels at a given rotor location defined between adjacent splitter and full blades at the given rotor location, a width of the first flow channels increasing from the hub to the tips of the splitter blades along at least a portion of the chord length of the splitter blades.

An aircraft defining a longitudinal direction and a lateral direction is provided. The aircraft includes: a fuselage; an engine mounted at a location spaced from the fuselage of the aircraft, the engine comprising a plurality of rotor blades; and a fuselage shield attached to or formed integrally with the fuselage at a location in alignment with the plurality of rotor blades along the lateral direction, the fuselage shield comprising a first layer defining a first density and a second layer defining a second density, the first density being different than the second density.

There is provided apparatuses and methods for a gas turbine engine. The embodiments include a core section with a flow path. The flow path includes a stator vane array having outlet vanes. Each outlet vane has a trailing edge. A strut has a leading edge that is upstream of the trailing edges. Alternatively, the flow path includes a stator vane array having inlet vanes. Each inlet vane has a leading edge. The strut has a trailing edge that is downstream of the leading edges. There also is increased spacing between adjacent vanes and rotor blades.

A tandem rotor disk apparatus may include a rotor disk body concentric about an axis. The tandem rotor disk apparatus may also include a first blade extending radially outward of the rotor disk body and a second blade extending radially outward of the rotor disk body. The first blade may be offset from the second blade in a direction parallel to the axis. The tandem rotor disk apparatus may be implemented in a gas turbine engine with no intervening stator vane stages disposed between the first blade and the second blade. The tandem rotor disk apparatus may include two separate rotor disk bodies.