A turbine blade and a radial turbine having at least one blade is provided. The turbine blade includes a trailing edge and a leading edge opposite the trailing edge. The turbine blade also includes a cooling passage defined internally within the turbine blade. The cooling passage is in fluid communication with a source of cooling fluid via a single inlet to receive a cooling fluid. The cooling passage diverges at a first point downstream from the single inlet into at least two branches that extend along the at least one blade from the first point to a second point near a tip of the leading edge and the cooling passage converges at the second point.

An airfoil for a gas turbine engine includes a first portion joined to a second portion along an interface such that at least the first portion establishes an airfoil section and the second portion establishes a root section. The airfoil section includes an airfoil body that extends between leading and trailing edges in a chordwise direction, between pressure and suction sides separated in a thickness direction, and from the root section to a tip portion in a spanwise direction. A recessed region extends inwardly from at least one of the pressure and suction sides. The airfoil body includes at least one rib that bounds a respective pocket within a perimeter of the recessed region. The recessed region and the at least one rib are dimensioned to extend across the interface. A cover skin is coupled to the airfoil body along the at least one rib to enclose the recessed region.

A fan includes a hub and a plurality of fan blades. The hub has an axle center. The fan blades are disposed around the hub. Each of the fan blades has a bent portion, and the bent portions of the fan blades are extended along a surrounding direction surrounding the axle center. The hub is welded with the bent portion of each of the fan blades along the surrounding direction. As a result, the number of fan blades is maximized, the strength is simultaneously ensured to be enough, and the advantages of effectively enhancing the fan characteristics are achieved.

Vane assemblies are described. The vane assemblies include a platform, an airfoil extending from the platform, a forward rail extending from the platform and arranged along a forward side of the platform, and an aft rail extending from the platform and arranged along an aft side of the platform. At least one support beam is provided extending in a forward-aft direction between the forward rail and the aft rail and separated from the platform by a first distance. The at least one support beam has a thickness in a radial direction of 40% or less of a total radial extent from the platform to an outer diameter edge of at least one of the forward rail and the aft rail and the at least one support beam has a thickness in a circumferential direction of 30% or less of a total circumferential extent of vane assembly.

A fan includes a hub and a plurality of fan blades. The hub has an axle center. The fan blades are disposed around the hub. Each of the fan blades has a bent portion, and the bent portions of the fan blades are extended along a surrounding direction surrounding the axle center. The hub is welded with the bent portion of each of the fan blades along the surrounding direction. As a result, the number of fan blades is maximized, the strength is simultaneously ensured to be enough, and the advantages of effectively enhancing the fan characteristics are achieved.

An airfoil for a gas turbine engine includes a first portion joined to a second portion along an interface such that at least the first portion establishes an airfoil section and the second portion establishes a root section. The airfoil section includes an airfoil body that extends between leading and trailing edges in a chordwise direction, between pressure and suction sides separated in a thickness direction, and from the root section to a tip portion in a spanwise direction. A recessed region extends inwardly from at least one of the pressure and suction sides. The airfoil body includes at least one rib that bounds a respective pocket within a perimeter of the recessed region. The recessed region and the at least one rib are dimensioned to extend across the interface. A cover skin is coupled to the airfoil body along the at least one rib to enclose the recessed region.

A method for manufacturing a turbo fan unit includes a step for preparing multiple fan blades and an other-side side plate, and a step for connecting each of the multiple fan blades to the other-side side plate by a welding process. In the preparing step, one of the fan blade and the side plate is prepared, in which a connecting-surface forming portion having a connecting surface and a welding projection protruded from the connecting surface is formed. The connecting surface connects one of the fan blade and the side plate to the other one of the fan blade and the side plate. In the connecting step, the welding projection is melted down and the connecting surface is connected to an opposing surface, which is the surface of one of the fan blade and the side plate and which is opposing to the connecting surface.

A fan includes a hub and a plurality of fan blades. The hub has an axle center. The fan blades are disposed around the hub. Each of the fan blades has a bent portion, and the bent portions of the fan blades are extended along a surrounding direction surrounding the axle center. The hub is welded with the bent portion of each of the fan blades along the surrounding direction. As a result, the number of fan blades is maximized, the strength is simultaneously ensured to be enough, and the advantages of effectively enhancing the fan characteristics are achieved.

A turbine blade is secured with a turbine disk by a retention assembly. The turbine blade includes a blade stem and the disk includes at least one mount post for mounting of the turbine blade. The retention assembly is secured with the disk by forming a diffusion joint with the disk.

A fan section for a gas turbine engine is disclosed herein. The fan may include a rotor disk and a plurality of airfoils fixedly attached to and supported by the rotor disk as a single unitary piece. The airfoils may extend radially outward from the rotor disk with respect to an engine axis. The rotor disk may be made of metal and the airfoils may each be made at least partially of an organic matrix composite.