A compressor includes a purge flow extraction path extending radially and configured to direct an airflow radially inwardly. Also included is a center bore at least partially defined by a rotor structure extending axially and fluidly coupled to the purge flow extraction path. Further included is an airflow manipulation device disposed entirely within the center bore, the airflow manipulation device having a plurality of vanes defining at least one vane slot.

A cooling structure for a turbine is configured such that: a plurality of disks rotating integrally with blades are arranged along a rotational axis; and the disks have formed therein disk holes arranged in a circumferential direction to supply cooling air for cooling the rotor blades to downstream disks. At least one of the disk holes is set such that, when the rotational direction of the disk is defined as the positive direction and the direction opposite the rotational direction is defined as the negative direction, an outlet absolute circumferential velocity vector which is a component in a rotational direction U of the velocity vector of the cooling air at an outlet of a disk hole is smaller than an inlet absolute circumferential velocity vector which is a component in the rotational direction of the velocity vector of the cooling air at an inlet of the disk hole.

A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk, a plurality of rotor blades and a plurality of disk mounts. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is mounted to the first rotor disk and to the second rotor disk. The rotor blades include a first rotor blade. Each of the disk mounts connects the first rotor disk and the second rotor disk together. The disk mounts include a first disk mount that further supports the first rotor blade.

A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk, a plurality of rotor blades and a plurality of disk mounts. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is axially between and mounted to the first rotor disk and the second rotor disk. The disk mounts connect the first rotor disk and the second rotor disk together. The disk mounts include a first disk mount. The first disk mount is integral with the first rotor disk. The first disk mount projects axially through the second rotor disk.

A rotor stack for a gas turbine engine includes a first rotor disk with a first rotor spacer arm, the first rotor spacer arm having a first flange with an outboard flange surface and an inboard flange surface, a first hole along an axis through the first flange, the first hole having a counterbore in the outboard flange surface; a second rotor disk with a web having a second hole along the axis; a third rotor disk with a third rotor spacer arm, the third rotor spacer arm having a third flange with an outboard flange surface and an inboard flange surface, a third hole along the axis through the third flange, the third hole having a counterbore in the inboard flange surface; and a bushing with a tubular body and a flange that extends therefrom, the tubular body comprising at least one axial groove along an outer diameter thereof, the bushing extends through the first hole, the second hole and partially into the counterbore in the inboard flange surface of the third hole.

A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk, a plurality of rotor blades and a plurality of vanes. The first rotor disk is configured to rotate about a rotational axis. The first rotor disk is configured from or otherwise includes disk material. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is axially between and mounted to the first rotor disk and the second rotor disk. The vanes are arranged circumferentially around the rotational axis and axially between the first rotor disk and the second rotor disk. The vanes include a first vane, which first vane is configured from or otherwise includes vane material that is different than the disk material.

A rotor for a gas turbine engine has: a first rotor disk; an interstage flange that extends from the first rotor disk to a flange end portion that has an axial end surface and first radial outer and inner surfaces; a circumferential groove, formed in the flange end portion and extending from the axial end surface toward the first rotor disk; radial outer and inner slots formed in the first radial outer and inner surfaces along the circumferential groove and extend through the first radial outer and inner surfaces; and a valve member disposed within the circumferential groove and is secured within the circumferential groove when the flange end portion is connected to a second rotor disk, wherein the valve member deflects from rotor rotational speeds to seal or unseal the radial outer slot.

A rotor disc for a turbomachine, the disc extending circumferentially about an axis and including a plurality of cavities configured to receive blade roots, each cavity including a downstream radial wall configured to axially block the blade root in the cavity, each downstream radial wall including a channel of ventilation of the cavity, including an inlet orifice which opens into the cavity and an outlet orifice which opens onto a downstream surface of the disc. An assembly for a turbomachine including such a disc and an upstream retention ring and a turbomachine including such an assembly.

A rotor disc for a turbomachine, the disc extending circumferentially around an axis and including a plurality of cells configured to receive blade roots and each cell including an upstream radial wall configured to axially block the blade root in the cell, each cell being connected to an upstream surface of the disc by a ventilation channel of the cell, the ventilation channel including an inlet orifice which opens onto the upstream surface of the disc and an outlet orifice which opens into the cell. An assembly of such a disc, of a plurality of blades and of a downstream retaining ring and a turbomachine including such an assembly.

Disclosed is an aero-engine turbine assembly, including a turbine assembly body and a cooling component. The turbine assembly body is provided with an internal flow passage, and the turbine assembly body includes a turbine rotor disk, a blade end wall and a turbine rotor blade, which are successively fixedly connected with each other. The internal flow passage passes through the turbine rotor disk, the blade end wall and the turbine motor blade, and the internal flow passage is provided with an inlet and an outlet. The cooling component is fixed on the turbine rotor disk, and the cooling component includes an electromagnetic pump system, an expansion joint and a radiator, which are successively communicated with each other. The electromagnetic pump system is communicated with the inlet, to inject liquid metal to into the internal flow passage.