A pressurized air supply system supplies, to a pressurization object device, flowing air that includes at least one of compressed air, which is generated by compressing air supplied from an air supply source, or discharged air from a turbocharger compressor forming a turbocharger. The compressor is controlled such that a saturated steam temperature of the flowing air supplied from the air supply source to the pressurization object device is lower than a temperature in the pressurization object device, at startup.

A pressurized air supply system supplies, to a pressurization object device, flowing air that includes at least one of compressed air, which is generated by compressing air supplied from an air supply source, or discharged air from a turbocharger compressor forming a turbocharger. The compressor is controlled such that a saturated steam temperature of the flowing air supplied from the air supply source to the pressurization object device is lower than a temperature in the pressurization object device, at startup.

An internal combustion engine, ICE, system, includes a turbocharger having a turbine and a compressor for compressing intake air and feeding the intake air to the ICE. A turbo turbine unit is disposed in an exhaust gas path downstream the turbocharger to receive exhaust gas from the turbocharger. The turbo turbine unit having a turbine wheel, a bearing housing defining an inside volume for containing lubrication oil, and a sealing arrangement positioned in the vicinity of the turbine wheel for preventing lubrication oil from escaping from the inside volume of the bearing housing to an exhaust duct of the turbo turbine unit. A buffer air conduit extends between a selected position at the compressor and a buffer air inlet of the turbo turbine unit. The buffer air inlet is in fluid communication with a buffer air channel inside the turbo turbine unit to direct buffer air to a position in-between axially opposite arranged annular sealing members to provide a counter-pressure against at least one of the annular sealing members.

A turbine engine having a compressor section, a combustor section, a turbine section, and a rotatable drive shaft that couples a portion of the turbine section and a portion of the compressor section. A bypass conduit couples the compressor section to the turbine section while bypassing at least the combustion section. At least one particle separator is located in the turbine engine having a separator inlet that receives a bypass stream, a separator outlet that receives a reduced-particle stream flows, and a particle outlet that receives a concentrated-particle stream comprising separated particles. A conduit, fluidly coupled to the particle outlet, extends through an interior of at least one stationary vane.

A turbine engine having a compressor section, a combustor section, a turbine section, and a rotatable drive shaft that couples a portion of the turbine section and a portion of the compressor section. A bypass conduit couples the compressor section to the turbine section while bypassing at least the combustion section. At least one particle separator is located in the turbine engine having a separator inlet that receives a bypass stream, a separator outlet that receives a reduced-particle stream flows, and a particle outlet that receives a concentrated-particle stream comprising separated particles. A conduit, fluidly coupled to the particle outlet, extends through an interior of at least one stationary vane.

A system includes an air cooling system having a heat exchanger, a fan, and a mount. The heat exchanger includes an inlet, an outlet, and a heat exchange conduit between the inlet and the outlet. The inlet is configured to couple to a bleed system of a gas turbine system to extract a bleed flow. The heat exchanger is configured to cool the bleed flow along the heat exchange conduit in a surrounding air to produce a cooled bleed flow. The outlet is configured to couple to a fuel purge system of the gas turbine system to supply the cooled bleed flow as a fuel purge flow. The fan is configured to force an airflow from the surrounding air through the heat exchanger. The mount is configured to mount the air cooling system outside of an enclosure surrounding the gas turbine system.

A system includes an air cooling system having a heat exchanger, a fan, and a mount. The heat exchanger includes an inlet, an outlet, and a heat exchange conduit between the inlet and the outlet. The inlet is configured to couple to a bleed system of a gas turbine system to extract a bleed flow. The heat exchanger is configured to cool the bleed flow along the heat exchange conduit in a surrounding air to produce a cooled bleed flow. The outlet is configured to couple to a fuel purge system of the gas turbine system to supply the cooled bleed flow as a fuel purge flow. The fan is configured to force an airflow from the surrounding air through the heat exchanger. The mount is configured to mount the air cooling system outside of an enclosure surrounding the gas turbine system.

A gas turbine engine includes a turbomachine, the turbomachine defining a core flow therethrough during operation. A first heat exchange assembly is in fluid communication with the turbomachine for receiving a first bleed flow from the turbomachine. A second heat exchange assembly is in fluid communication with the turbomachine for receiving a second bleed flow from the turbomachine. A first flow outlet is provided for receiving the first bleed flow from the first heat exchange assembly and providing the first bleed flow to a first aircraft flow assembly. A second flow outlet is provided for receiving the second bleed flow and providing the second bleed flow from the second heat exchange assembly to a second aircraft flow assembly.

A gas turbine engine includes a turbomachine, the turbomachine defining a core flow therethrough during operation. A first heat exchange assembly is in fluid communication with the turbomachine for receiving a first bleed flow from the turbomachine. A second heat exchange assembly is in fluid communication with the turbomachine for receiving a second bleed flow from the turbomachine. A first flow outlet is provided for receiving the first bleed flow from the first heat exchange assembly and providing the first bleed flow to a first aircraft flow assembly. A second flow outlet is provided for receiving the second bleed flow and providing the second bleed flow from the second heat exchange assembly to a second aircraft flow assembly.

A gas turbine engine comprises a turbomachine defining a core flow therethrough during operation. A flow tap is provided in fluid communication with the turbomachine, wherein the flow tap is configured to receive a portion of the core flow therethrough as a bleed flow. A bleed assembly includes a machine load, a bleed flow machine, and a bleed regulator. The bleed flow machine is disposed in fluid communication with the turbomachine through the flow tap, and is configured to drive the machine load. The bleed regulator is configured to regulate a bleed output provided to the bleed flow machine by controlling a capture rate of the bleed flow by the bleed flow machine.