The leading edge, the trailing edge, or both may be axially offset for a portion of the airfoils in a disk. By offsetting the airfoils, the downstream wake energy to the next stage of airfoils may be decreased. By staggering airfoils which are offset with airfoils that are not offset, the wake shapes from the airfoils may be out of phase and will not excite the downstream airfoils as much as conventional systems. This may decrease vibration and associated vibratory stresses in the system.

Provided is a turbine airfoil including: a cooling passage that allows a cooling medium to move from a base part side to a tip end part side in an airfoil height direction; a lattice structure including rib sets stacked in a lattice pattern in the cooling passage; inverting portions at opposite side edge portions of the lattice structure, each being open at a side edge portion and allowing the cooling medium to be inverted from a lattice flow passage defined between ribs of one rib set to a lattice flow passage defined between ribs of another rib set; and a communication flow passage defined between one side edge portion of the lattice structure and a side wall surface of the cooling passage, the communication flow passage extending in the airfoil height direction to communicate a plurality of lattice flow passages at the one side edge portion.

A method for manufacturing a blade in composite material having an added metal leading edge for gas turbine aeroengine, includes producing a blade body in composite material including in longitudinal direction, a blade root part, a shank part and an airfoil body part; manufacturing, via additive manufacturing, a leading edge extending in longitudinal direction between a lower end present at the shank part in composite material and an upper end present at the tip; bonding the manufactured leading edge onto the foremost edge portion of the airfoil body of the blade body in composite material.

A rotor blade system includes a rotor, a casing radially spaced apart from the rotor, and a plurality of blades coupled to the rotor and positioned between the rotor and the casing. The one or more of the plurality of blades have a radial length different from a remaining one or more of the plurality of blades so as to vary a tip gap between a tip of the one or more of the plurality of blades and the casing to break up a frequency content of a leakage vortex at the tip to modify natural frequencies of the plurality of blades and mode shapes to reduce or substantially eliminate flutter.

A blade for a turbo machine is provided. The blade for a turbo machine includes an airfoil body extending in a radial direction and including a suction side surface and a pressure side surface opposite to the suction side surface with respect to a circumferential direction extending across the radial direction, and a snubber structure including a first snubber element protruding in the circumferential direction from the suction side surface of the airfoil body and a second snubber element protruding in the circumferential direction from the pressure side surface of the airfoil body, wherein the first snubber element is connected to the suction side surface of the airfoil body by a concave curved first transition portion having a first radius, and the second snubber element is connected to the pressure side surface of the airfoil body by a concave curved second transition portion having a second radius, the first radius being smaller than the second radius.

Various embodiments of the disclosure include turbine blades and systems employing such blades. Various embodiments include a turbine blade having: an airfoil having an airfoil shape having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in TABLE I. The Cartesian coordinate values are non-dimensional values of from 0% to 100% convertible to distances by multiplying the values by a height of the airfoil expressed in units of distance. The X and Y values are connected by smooth continuing arcs to define airfoil profile sections at each distance Z along at least a portion of the airfoil, the profile sections at the Z distances being joined smoothly with one another to form the nominal profile.

The present invention relates to a rotor blade (20) for arrangement in a gas duct (2) of a turbomachine (1), having a rotor blade airfoil (23), which, viewed in a tangential section, has a blade airfoil profile (24) with a leading edge radius RVK and a rotor blade airfoil thickness d, wherein the blade airfoil profile (24) is thickened, at least in sections, specifically the blade airfoil thickness d is specified, in relation to the front edge radius RVK, such that (2d/Rvk2)−d≤5.5.

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 blade and gas turbine include a stationary blade main body provided internally with cavities, and an inner shroud linked to an end portion, in the longitudinal direction, of the stationary blade main body, and which is internally provided with an inner shroud cooling passage with which a first cavity is in fluid communication. The inner shroud is provided, in front edge corner portions, with a first chamfered portion intersecting a front surface, a side surface, and an upper surface, and first cooling holes in fluid communication with the inner shroud cooling passage are provided in the first chamfered portion.

A rotary machine is provided. The rotary machine includes a rotor disk including a plurality of slots, a plurality of blades mounted on an outer circumferential surface of the rotor disk, each of the blades having a root part inserted into an associated one of the slots, and a retainer restraining axial movement of the plurality of blades, wherein the retainer includes an inner fixing member inserted into a bottom of the slot, an outer fixing member inserted into a lower surface of the root part to abut against the inner fixing member, and a spacer inserted between the outer fixing member and the bottom of the slot to prevent the outer fixing member from being separated.