Abstract: A thermal management system for a gas turbine engine includes a high pressure compressor rotor having a plurality of disks and a shaft; at least one forward cavity between a rim of at least one of the disks and the shaft and at least one aft cavity formed between a rim of at least one other of the disks and the shaft; a first flow of cooling air in the at least one forward cavity; and a second flow of cooling air in the at least one aft cavity.

Abstract: A thermal management system for a gas turbine engine includes a high pressure compressor rotor having a plurality of disks and a shaft; at least one forward cavity formed between a rim of at least one of the disks and the shaft and at least one aft cavity formed between a rim of at least one other of the disks and the shaft; a first flow of cooling air in the at least one forward cavity; and a second flow of cooling air in the at least one aft cavity.

Abstract: A seal assembly includes a first component, a second component that defines an outer surface and is located radially inward from the first component, a groove defined in the second component and arranged to face the first component, and a seal ring positioned between the first and second components and extending at least partially into the groove. The seal ring defines an outer diameter surface and a first lateral surface adjoining the outer diameter surface. The seal ring is split to define a first free end and a second free end, and the first and second free ends are configured to overlap along a split surface that extends from the first lateral surface to the outer diameter surface.