An apparatus is provided that includes a rotor bore and a bore basket. The bore basket extends axially between a first and second axial ends. The bore basket includes an engagement panel, and cylindrical inner and outer radial panels. The inner and outer radial panels extend substantially between the first and second axial ends. The inner radial panel is disposed radially inside of and separated from the outer radial panel, defining an annular passage disposed there between. The bore basket includes a plurality of inlet apertures in fluid communication with the annular passage at a first axial position, and a plurality of exit apertures in fluid communication with the annular passage at a second axial position. The second axial end of the bore basket is attached to an inner radial hub of the rotor bore, and the rotor bore and bore basket are a unitary structure.

A blade track system includes a blade track configured to be positioned around a plurality of blades of a gas turbine engine. The blade track includes: a full-hoop body, an inlet port, and an exit port. The full-hoop body includes an outer surface and an inner surface. The inner surface defines an internal cooling channel extending within the full-hoop body. The inlet port and the exit port are each integrated with the full-hoop body and in fluid communication with the internal cooling channel.

A turbine engine having counter-rotating rotors comprising a first rotor, rotating in a first rotational direction, defining a first rotor set of blades axially spaced to define a gap, and a second rotor, rotating in a second rotational direction counter the first rotational direction. The second rotor further including a second set of blades received within the gap of the first rotor. A plurality of fluid passages is formed in the first rotor with an outlet facing the gap.

A turbine rotor of an embodiment includes a rotor body portion having turbine discs in plural stages in an axial direction protruding radially outside from an outer peripheral surface of the rotor body portion over a circumferential direction. The turbine rotor includes a plurality of axial passages, through which a cooling medium flows, formed at the rotor body portion radially outside than a center axis of the turbine rotor and radially inside than an outer peripheral surface of the rotor body portion in the axial direction; an introduction passage introducing the cooling medium into each of the axial passages; and discharge passages that penetrate from each of the axial passages to the outer peripheral surface of the rotor body portion to discharge the cooling medium.

A flow diverter for an air separator of a gas turbine includes a cylindrical body configured to fit within a cooling hole of the air separator. One or more air flow vents are defined through and around a partial circumference of the cylindrical body. A bottom panel closes the cylindrical body at one end. A mounting flange surrounds the cylindrical body at an open end and extends radially outward from the cylindrical body. When the flow diverter is installed, air flows through the open end in a radial direction and exits through the air flow vents in an axial direction. The cylindrical body may include a collapsible region that collapses to engage the air separator and prevent the flow diverter from being dislodged. The air separator may additionally or alternately include supplemental cooling holes in a recessed area proximate to its mounting flange.

A flow guiding appliance of a turbomachine, in particular of an aircraft engine, for a partial exit flow of an outlet guide vane of a compressor, wherein, a flow guiding element delimits an annular channel about a shaft of the turbomachine. A frictional engagement connecting surface at the circumference of the flow guiding element connects the flow guiding element to another structural component of the turbomachine. The frictional engagement connecting surface with a closed enveloping surface can be inserted in a deformed state into the structural component with a circular cylindrical sealing surface with at least two opposite points. The at least two points of the closed enveloping surface are arranged in at least two opposite frictionally engaged contact areas of the circular cylindrical sealing surface following deformation. The invention also relates to a method for creating a flow guiding element.

A gas turbine engine includes a combustion section and a compressor section, the compressor section including a high pressure compressor. The high pressure compressor includes an aft-most compressor stage and an upstream compressor stage, each of the aft-most compressor stage and the upstream compressor stage including a rotor disk. The gas turbine engine also includes a high pressure spool assembly, the high pressure spool assembly including a rotor disk, and an airflow member extending from the rotor disk of the high pressure spool assembly to the rotor disk of the upstream compressor stage of the high pressure compressor to define in part a compressor cooling air passage outward of the airflow member along a radial direction.

A rotor assembly of a gas turbine engine may be spoked and includes a rotor and a shell. The rotor has a rotor disk and a plurality of blades each having a platform attached to the rotor disk and with a first channel defined radially between the platforms and the rotor disk. The shell projects aft of the rotor and includes inner and outer walls with a passage defined therebetween. The passage is in fluid communication with the first channel and, together, form part of a secondary flowpath for cooling of adjacent components. The rotor assembly may further include a structure located radially inward of the rotor disk and shell. The structure defines a supply conduit for flowing air from the passage and into a rotor bore defined at least in part by adjacent rotor disks. The entering air, being pre-heated when flowing through the channel and passage, warms the bore and reduces thermal gradients, thus thermal fatigue, across the rotor disk.

A cooling system configured to cool aspects of the turbine engine between a compressor and a turbine assembly is disclosed. In at least one embodiment, the cooling system may include one or more mid-frame cooling channels extending from an inlet through one or more mid-frame torque discs positioned downstream of the compressor and upstream of the turbine assembly. The inlet may be positioned to receive compressor bleed air. The mid-frame cooling channel may be positioned in a radially outer portion of the mid-frame torque disc to provide cooling to outer aspects of the mid-frame torque disc such that conventional, low cost materials may be used to form the mid-frame torque disc rather than high cost materials with capacity to withstand higher temperatures. The cooling fluid routed through the mid-frame cooling channel in the mid-frame torque disc may be exhausted into a cooling system (10) for the downstream turbine assembly.

A compressor section or a turbine section of a gas turbine engine having an axis includes a drum. The compressor section or the turbine section also includes a plurality of bores extending radially inward from the drum including a first bore and a second bore. The compressor section or the turbine section also includes a first bore basket at least partially defining a first cavity such that the first bore has at least one surface located in the first cavity. The compressor section or the turbine section also includes a second bore basket at least partially defining a second cavity that is isolated from the first cavity such that the second bore has at least one surface located in the second cavity.