A system and method for improved system efficiency of an integrated power and control unit (IPCU) of an aircraft is disclosed. The system uses an open-loop cooling system and turbo machine power matching to provide wide operation range without over-sizing. In order to reduce the temperature of the air flow through the cooling heat exchanger, the cooling turbine need to expand further in the same time generating power but the power could be higher than the compressor could absorb so a generator that would convert the power and used in supplying the aircraft would result in more efficient system.

A gas turbine engine having: an engine core having, in serial flow communication, a low-pressure compressor, a high-pressure compressor, a high-pressure turbine drivingly connected to the high-pressure compressor, and a low-pressure turbine drivingly connected to an output shaft; and a clutch having a disengaged configuration in which the low-pressure turbine is drivingly disconnected from the low-pressure compressor such that, in the disengaged configuration, the clutch disengages the low-pressure turbine from the low-pressure compressor, and an engaged configuration in which the low-pressure turbine is drivingly connected to the low-pressure compressor, the low-pressure turbine drivingly connected to the output shaft in both of the engaged and disengaged configurations of the clutch.

Systems and techniques for engine starter control may include determining a current speed band associated with an engine start for a current time increment, determining a previous speed band associated with the engine start for a previous time increment within the engine start, receiving a measured current value, determining a target current value based on the current speed band, the previous speed band, whether an acceleration or deceleration occurred between the current speed band and the previous speed band, and a mode of a system for engine starter control, and adjusting a current for the engine start based on the target current value.

A system for starting a turbine engine. The system may comprise a gearbox, a first starter, and a second starter. The gearbox may have a gearbox input shaft. The gearbox may be coupled to the turbine engine. The gearbox input shaft may be rotatively coupled to a spool of the turbine engine. The first starter may be coupled to the gearbox input shaft. The second starter may have a second-starter output shaft. The second-starter output shaft may be coaxial with the gearbox input shaft. The second starter may be coupled to the gearbox input shaft through the first starter.

The invention relates to an architecture of a propulsion system of a multi-engine helicopter comprising turboshaft engines connected to a power transmission gearbox, characterised in that it comprises: at least one hybrid turboshaft engine (20) capable of operating in at least one standby mode during a stable cruise flight of the helicopter; at least two systems (30; 40) for controlling each hybrid turboshaft engine (20), each system (30; 40) comprising an electric machine (31; 41) connected to the hybrid turboshaft engine (20) and suitable for rotating the gas generator thereof, and at least one source (33; 43) of electrical power for said electric machine (31; 41), each reactivation system (30; 40) being configured such that it can drive said turboshaft engine (20) in at least one operating mode among a plurality of predetermined modes.

An aircraft engine including a fan having a center of mass, a low-pressure shaft that couples the fan directly to a low-pressure turbine of the aircraft engine, and a bearing arrangement for mounting the low-pressure shaft, where the bearing arrangement has at least two front bearings for mounting the low-pressure shaft, a first front bearing and a second front bearing, with the first front bearing being arranged in the axial direction in front of the second front bearing. It has been provided that the first front bearing is arranged substantially in the same plane, extending perpendicular to the longitudinal axis of the aircraft engine, as the center of mass of the fan.

A method of cooling a shaft of a gas turbine engine includes moving a control valve towards a position that inhibits fluid flow from a high pressure air source to an air turbine starter and enables fluid flow from a blower motor to the air turbine starter, in response to a gas turbine engine shutdown. The method further includes operating the blower motor to provide air to the air turbine starter.

The present disclosure is directed to an oil-free gas turbine engine. The gas turbine engine includes a compressor section, a combustion section, a turbine section, and an exhaust nozzle section. Further, the gas turbine engine includes at least one rotary component configured to drivingly connect at least a portion of the turbine section to at least a portion of the compressor section. Moreover, the gas turbine engine includes one or more gas-lubricated bearings configured to support the rotary component. In addition, the gas turbine engine includes a direct-drive starter-generator configured to start the gas turbine engine. Thus, the gas turbine engine of the present disclosure provides an engine that is at least partially oil free.

An electric machine (212) comprises a turbomachine rotor (203) having a hub (302) and an axis of rotation (A-A) about which the turbomachine rotor is arranged to rotate. The turbomachine rotor includes a plurality of blades (301). Each blade has a root (303) attached to the hub, a tip (304) remote from the hub, a leading edge (305) and a trailing edge (306), a pressure side and a suction side (307). A stator (502) is located circumferentially around the turbomachine rotor. Each blade further comprises a rotor element at the tip comprising a permanent magnet having a first pole (401) and a second pole (402), the first pole being located adjacent the suction side of the blade and the second pole being located adjacent the pressure side such that a magnetic flux path extends perpendicularly through the blade tip.

There is provided a rotor bow mitigation system and method for an aircraft engine. At least one value of at least one engine parameter prior to a shutdown of the engine is obtained, the at least one engine parameter comprising a first temperature internal to the engine. A second temperature external to the engine is measured and a motoring duration and a motoring interval for the engine are determined based on at least the first temperature and on the second temperature. Upon detecting a start indication for the engine, the engine is motored for the motoring duration and at the motoring interval.