A gaspath component includes a body having at least one internal cooling passage and a metering feature integral to the internal cooling passage. The metering feature and the body are a unitary structure.

A gas turbine engine includes a rotor carrying a blade, a ceramic vane adjacent the blade, a seal carried on the rotor adjacent the tip of the ceramic vane, and a cooling passage circuit extending through the blade, the seal, and the ceramic vane. The cooling passage circuit is configured to provide cooling air into the blade, then from the blade into the seal, and then from the seal into the ceramic vane.

A turbine blade of a turbine rotor blade ring, having a suction side, a pressure side and a cooling air duct through which a cooling medium is conveyable for cooling the turbine blade. It is provided that the cooling air duct has in at least one section a course such that its cross-sectional surface increases in the flow direction of the cooling medium up to a maximum in a first, widening partial section, its cross-sectional surface decreases in a second, narrowing partial section behind the maximum, and the cooling medium in the second, narrowing partial section is accelerated with a directional component in the direction of the suction side of the turbine blade. The invention furthermore relates to a method for conveying a cooling medium in a turbine blade of a turbine rotor blade ring.

A turbine blade having a base and an airfoil, the base including cooling air inlets and an internal cooling air passageway, and the airfoil including an internal multi-bend heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil. The airfoil also includes a “skin” that encompasses a tip wall, an inner spar, and a tip flag cooling system.

A turbine, a gas turbine including the same, and a method of assembling and disassembling the same, capable of ensuring stable sealing performance as well as facilitating assembly and disassembly are provided. The turbine may include a rotor disk having a lower hook, a plurality of turbine blades installed on an outer peripheral surface of the rotor disk, each of the turbine blades having a blade hook located above the lower hook, and a retainer sealing a cooling passage defined between the turbine blade and the rotor disk and supported by the blade hook and the lower hook.

A tangential on-board injector (TOBI) having: a body defining an annular passageway, discharge nozzles; a rotating component configured to be mounted for rotation relative to the body about an axis of rotation; a seal extending between the body and the rotating component; and flow passages circumferentially distributed about the axis of rotation and in fluid communication with the nozzles and the seal, the flow passages located upstream of the seal relative to a flow of the cooling air circulating toward the seal from the plurality of discharge nozzles, each of the flow passages extending along a respective passage axis, the passage axis of at least one of the flow passages having a tangential component at an outlet of the at least one of the flow passages that is different than a tangential component of an exit flow axis of at least one of the plurality of discharge nozzles.

A turbine assembly for use with a gas turbine engine includes a bladed wheel assembly, a vane assembly, and an inner seal. The bladed wheel assembly is adapted to interact with gases flowing through a gas path of the gas turbine engine. The vane assembly is located upstream of the bladed wheel assembly and adapted to direct the gases at the bladed wheel assembly. The inner seal is configured to block gases from passing around the vane assembly.

A control device controls a temperature of a shaft seal portion provided around a rotating shaft of a rotating machine to an appropriate temperature by adjusting an amount of the cooling air to be supplied to the shaft seal portion. The control device calculates a sensitivity indicated using the temperature of the shaft seal portion with respect to a flow rate of the cooling air supplied to the shaft seal portion and controls the flow rate of the cooling air so that the sensitivity has a predetermined target value on the basis of the calculated sensitivity. When the sensitivity is calculated, the flow rate is varied in a predetermined range having a certain flow rate as a center. The sensitivity of the temperature at the flow rate as a center is calculated from the variation in the temperature of the shaft seal portion with respect to the variation in the flow rate.

An on-board injector that delivers discharge air toward a turbine rotor of a gas turbine engine includes a second wall spaced form a first wall to define an annular inlet about an engine longitudinal axis and a multiple of airfoil shapes between the first wall and the second wall to segregate discharge air from the annular inlet, and a multiple of bypass apertures each along a radial axis transverse to the engine longitudinal axis through each of the multiple of airfoil shapes and the respective first wall, the second wall.

The invention relates to an assembly (1) for a turbine of a turbomachine, comprising: a first rotor disk (20a), a second rotor disk (20b), a part forming a mobile ring (28), comprising a system for preventing rotation of the mobile ring (28) relative to the rotor disks (20a, 20b), said system comprising: a rotor disk securing flange (222) having a plurality of teeth (224) that are distributed circumferentially about the turbomachine longitudinal axis (X-X), and a mobile ring securing flange (282) having a plurality of lugs (284) that are distributed circumferentially about the turbomachine longitudinal axis (X-X), the engagement of the rotor disk securing flange (222) with the mobile ring securing flange (282) ensuring, by means of the teeth (224) and the lugs (284), that the mobile ring (28) does not rotate relative to the rotor disks (20a, 20b).