Abstract: An aircraft propulsion system includes at least first and second electrical generators, each being configured to provide electrical power to a respective first and second AC electrical network. The system further comprises at least first and second AC electrical motors directly electrically coupled to a respective AC network and coupled to a respective propulsor, and a DC electrical network electrically coupled to the first and second AC networks via respective first and second AC to DC converters, and to a further electrical motor, the further electrical motor being coupled to a propulsor.

Abstract: An aircraft propulsion system includes at least first and second electrical generators, each being configured to provide electrical power to a respective first and second AC electrical network. The system further comprises at least first and second AC electrical motors directly electrically coupled to a respective AC network and coupled to a respective propulsor, and a DC electrical network electrically coupled to the first and second AC networks via respective first and second AC to DC converters, and to a further electrical motor, the further electrical motor being coupled to a propulsor.

Abstract: An aircraft propulsion system (10) comprises at least first and second electrical generators (15a, 15b), each being configured to provide electrical power to a respective first and second AC electrical network (16a, 16b). The system (10) further comprises at least first and second AC electrical motors (19a, 19b) directly electrically coupled to a respective AC network (16a, 16b) and coupled to a respective propulsor (4), and a DC electrical network electrically coupled to the first and second AC networks (16a, 16b) via respective first and second AC to DC converters (17a, 17b), and to a further electrical motor 19c), the further electrical motor (19c) being coupled to a propulsor (4).

Abstract: A gas turbine engine (10) having a bearing support structure (100) comprising: a bearing chamber (104); a first buffer fluid passage (110) disposed radially outwardly of the bearing chamber (104); a second buffer fluid passage (116) disposed radially outwardly of the first buffer fluid passage (110); and a core fluid passage (132) disposed radially outwardly of the second buffer fluid passage (116); the bearing chamber (104), first buffer fluid passage (110), second buffer fluid passage (116) and core fluid passage (132) being separated by respective hub walls.

Abstract: A gas turbine engine includes a spool which rotates in use, a static structure and an air bearing, the air bearing being provided at an interface between the spool and the static structure.

Abstract: An aircraft pneumatic system includes a main gas turbine engine including a main compressor, a recuperated auxiliary gas turbine engine including a core compressor having an inlet in fluid communication with a main compressor bleed of the main gas turbine engine and an environmental control system inlet in fluid communication with the main compressor bleed of the main gas turbine engine. The environmental control system includes a compressor driveable by power provided by the auxiliary gas turbine engine.

Abstract: An aircraft propulsion system (10) comprises at least first and second electrical generators (15a, 15b), each being configured to provide electrical power to a respective first and second AC electrical network (16a, 16b). The system (10) further comprises at least first and second AC electrical motors (19a, 19b) directly electrically coupled to a respective AC network (16a, 16b) and coupled to a respective propulsor (4), and a DC electrical network electrically coupled to the first and second AC networks (16a, 16b) via respective first and second AC to DC converters (17a, 17b), and to a further electrical motor 19c), the further electrical motor (19c) being coupled to a propulsor (4).

Abstract: An aircraft pneumatic system comprises a main gas turbine engine (10) comprising a main compressor (14, 16), a recuperated auxiliary gas turbine engine (40) comprising a core compressor (42) having an inlet in fluid communication with a main compressor bleed (32) of the main gas turbine engine(10) and an environmental control system inlet in fluid communication with the main compressor bleed (32) of the main gas turbine engine (10). The environmental control system comprises a compressor (68) driveable by power provided by the auxiliary gas turbine engine (40).

Abstract: A gas turbine engine includes a spool which rotates in use, a static structure and an air bearing, the air bearing being provided at an interface between the spool and the static structure.

Abstract: A cooling arrangement for a compressor drive cone (48) in a gas turbine engine (10). An annular cavity (62) defined by the drive cone (48), a portion of a compressor rotor disc (44) and a downstream cavity wall (52). The cavity (62) comprising a recirculation zone (66) at its radially outermost region. The cavity (62) arranged to ingest hot working gas radially inwardly into the recirculation zone (66). An annular array of pipes (70) that penetrate through the cavity wall (52) close to the radially inward extent of the recirculation zone (66). The pipes (70) coupled to a supply of cooling fluid and having an open end arranged to direct the cooling air, in use, across the recirculation zone (66) towards the drive cone (48).