Abstract: A compressor for use in a gas turbine engine comprises a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A radially outer housing surrounds an outer diameter of the blades. A lower pressure tap and a higher pressure tap tap air from two distinct locations within the compressor and radially outwardly through the outer housing. A valve selectively delivers at least one of the lower pressure tap and the higher pressure tap to the bore of the disc. A control for the valve is programmed to move the valve to a position delivering the higher pressure tap at a point prior to take-off when the compressor is mounted in a gas turbine engine on an aircraft. A gas turbine engine and a method of operating a gas turbine engine are also disclosed.

Abstract: A method of monitoring a gas turbine engine includes the steps of: (a) receiving information from actual flights of an aircraft including an engine to be monitored, and including at least one of the ambient temperature at takeoff, and internal engine pressures, temperatures and speeds; (b) evaluating the damage accumulated on an engine component given the data received in step (a); (c) storing the determined damage from step (b); (d) repeating steps (a)-(c); (e) recommending a suggested future use for the component based upon steps (a)-(d). A system is also disclosed.

Abstract: A method of monitoring a gas turbine engine includes the steps of: (a) receiving information from actual flights of an aircraft including an engine to be monitored, and including at least one of the ambient temperature at takeoff, and internal engine pressures, temperatures and speeds; (b) evaluating the damage accumulated on an engine component given the data received in step (a); (c) storing the determined damage from step (b); (d) repeating steps (a)-(c); (e) recommending a suggested future use for the component based upon steps (a)-(d). A system is also disclosed.

Abstract: A compressor section for use in a gas turbine engine is comprised of a plurality of compressor stages, with each stage including a disk having an inner periphery defining a bore that faces a shaft. A flow path flows in a generally axial direction between the shaft and the bores of each disk. At least one deflector is positioned between a pair of adjacent bores to direct the air flow radially outwardly into a cavity formed between an associated pair of adjacent disks.

Abstract: A method of monitoring a gas turbine engine includes the steps of: (a) receiving information from actual flights of an aircraft including an engine to be monitored, and including at least one of the ambient temperature at takeoff, and internal engine pressures, temperatures and speeds; (b) evaluating the damage accumulated on an engine component given the data received in step (a); (c) storing the determined damage from step (b); (d) repeating steps (a)-(c); (e) recommending a suggested future use for the component based upon steps (a)-(d). A system is also disclosed.

Abstract: A rotor disk for a turbomachine includes a disk body, a rim configured to connect to or include a rotor blade disposed on a radially outward portion of the disk body, a bore defined in a radially inward portion of the disk body and configured to be radially adjacent to a shaft, and a thermal regulation aperture defined in the disk body, radially inward of the rim, for allowing flow to pass axially through the disk body when disposed in a stage of a turbomachine, and a thermal regulation aperture defined in the disk body for allowing flow to pass through the disk body when disposed in a stage of a turbomachine.

Abstract: A compressor for use in a gas turbine engine comprises a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A high pressure air tap includes a lower temperature tapped path and a higher temperature tapped path and a valve for selectively delivering one of the lower temperature tapped path and the higher temperature tapped path into the bore of the disc. The valve is operable to selectively block flow of either of the lower pressure and higher pressure tapped paths to the bore of the disc, with the disc including holes to allow air from compressor chambers to communicate with the bore of the disc. A gas turbine engine and a method of operating a gas turbine engine are also disclosed.

Abstract: A method of monitoring a gas turbine engine includes the steps of: (a) receiving information from actual flights of an aircraft including an engine to be monitored, and including at least one of the ambient temperature at takeoff, and internal engine pressures, temperatures and speeds; (b) evaluating the damage accumulated on an engine component given the data received in step (a); (c) storing the determined damage from step (b); (d) repeating steps (a)-(c); (e) recommending a suggested future use for the component based upon steps (a)-(d). A system is also disclosed.

Abstract: A gas turbine engine comprises a compressor including a disk and a blade. A turbine rotor has a disk and a blade. Turbine conditioning air is supplied to the turbine rotor. The turbine conditioning air passes across the disk of the compressor to condition the disk. A method of operating a gas turbine engine is also disclosed.

Abstract: A compressor section for use in a gas turbine engine is comprised of a plurality of compressor stages, with each stage including a disk having an inner periphery defining a bore that faces a shaft. A flow path flows in a generally axial direction between the shaft and the bores of each disk. At least one deflector is positioned between a pair of adjacent bores to direct the air flow radially outwardly into a cavity formed between an associated pair of adjacent disks.

Abstract: A compressor for use in a gas turbine engine comprises a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A high pressure air tap includes a lower temperature tapped path and a higher temperature tapped path and a valve for selectively delivering one of the lower temperature tapped path and the higher temperature tapped path into the bore of the disc. The valve is operable to selectively block flow of either of the lower pressure and higher pressure tapped paths to the bore of the disc, with the disc including holes to allow air from compressor chambers to communicate with the bore of the disc. A gas turbine engine and a method of operating a gas turbine engine are also disclosed.

Abstract: A compressor for use in a gas turbine engine comprises a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A radially outer housing surrounds an outer diameter of the blades. A lower pressure tap and a higher pressure tap tap air from two distinct locations within the compressor and radially outwardly through the outer housing. A valve selectively delivers at least one of the lower pressure tap and the higher pressure tap to the bore of the disc. A control for the valve is programmed to move the valve to a position delivering the higher pressure tap at a point prior to take-off when the compressor is mounted in a gas turbine engine on an aircraft. A gas turbine engine and a method of operating a gas turbine engine are also disclosed.

Abstract: A turbomachine includes an aft stage including a rotor disk having a rotor disk bore hole and a rotor disk rim, and a valve disposed within the turbomachine and in fluid communication with the rotor disk bore hole. The valve is in selective fluid communication with a first pressure source such that, in an first position, the valve allows fluid to flow from the first pressure source through the rotor disk bore hole and radially outward along the aft rotor disk toward the aft rotor disk rim.

Abstract: A rotor disk for a turbomachine includes a disk body, a rim configured to connect to or include a rotor blade disposed on a radially outward portion of the disk body, a bore defined in a radially inward portion of the disk body and configured to be radially adjacent to a shaft, and a thermal regulation aperture defined in the disk body, radially inward of the rim, for allowing flow to pass axially through the disk body when disposed in a stage of a turbomachine, and a thermal regulation aperture defined in the disk body for allowing flow to pass through the disk body when disposed in a stage of a turbomachine.