A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.

A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.

An aspect includes a method for motoring control for multiple engines of an aircraft is provided. A controller can determine a motoring time of a first engine starting system to cool a first engine. The controller can compare the motoring time of the first engine starting system with a motoring time of one or more other engine starting systems of one or more other engines of the aircraft. The motoring time of the first engine starting system can be controlled relative to a tolerance of the motoring time of the one or more other engine starting systems by adjusting the motoring time of the first engine starting system relative to the one or more other engine starting systems in a motoring sequence based on comparing the motoring time of the first engine starting system with the motoring time of the one or more other engine starting systems.

An aspect includes a method for motoring control for multiple engines of an aircraft is provided. A controller can determine a motoring time of a first engine starting system to cool a first engine. The controller can compare the motoring time of the first engine starting system with a motoring time of one or more other engine starting systems of one or more other engines of the aircraft. The motoring time of the first engine starting system can be controlled relative to a tolerance of the motoring time of the one or more other engine starting systems by adjusting the motoring time of the first engine starting system relative to the one or more other engine starting systems in a motoring sequence based on comparing the motoring time of the first engine starting system with the motoring time of the one or more other engine starting systems.

A through-flow gas turbine engine includes a core comprising multiple spools rotatable about a center axis. An accessory gearbox (AGB) is drivingly engaged to the core and disposed aft of the outlet. A reduction gearbox (RGB) is drivingly engaged to the core and disposed forward of the inlet. The RGB has an RGB output to provide rotational output to a rotatable load. An electric motor is drivingly engaged to the rotatable load. An electric generator is configured to provide electrical power to the electric motor. One of the electric motor and the electric generator is disposed axially between the outlet and the AGB and the other of the electric motor and the electric generator is disposed axially between the inlet and the RGB.

A through-flow gas turbine engine includes a core comprising multiple spools rotatable about a center axis. An accessory gearbox (AGB) is drivingly engaged to the core and disposed aft of the outlet. A reduction gearbox (RGB) is drivingly engaged to the core and disposed forward of the inlet. The RGB has an RGB output to provide rotational output to a rotatable load. An electric motor is drivingly engaged to the rotatable load. An electric generator is configured to provide electrical power to the electric motor. One of the electric motor and the electric generator is disposed axially between the outlet and the AGB and the other of the electric motor and the electric generator is disposed axially between the inlet and the RGB.

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

A hybrid power generation unit and synchronous condenser system connectable to a power grid includes a combustion turbine coupled to a first shaft and operable to provide rotational energy to the first shaft, a gear box coupled to the first shaft, and a first clutch portion coupled to the first shaft. A motor is selectively coupled to the gear box to turn the gear box and the first shaft, a second clutch portion is connected to a second shaft, and a generator is coupled to the second shaft. The generator is selectively connectable to the grid to operate as a synchronous condenser when the first clutch portion and the second clutch portion are disengaged and to convert rotational energy from the first shaft to electrical power when the first clutch portion and the second clutch portion are engaged.

A hybrid power generation unit and synchronous condenser system connectable to a power grid includes a combustion turbine coupled to a first shaft and operable to provide rotational energy to the first shaft, a gear box coupled to the first shaft, and a first clutch portion coupled to the first shaft. A motor is selectively coupled to the gear box to turn the gear box and the first shaft, a second clutch portion is connected to a second shaft, and a generator is coupled to the second shaft. The generator is selectively connectable to the grid to operate as a synchronous condenser when the first clutch portion and the second clutch portion are disengaged and to convert rotational energy from the first shaft to electrical power when the first clutch portion and the second clutch portion are engaged.

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.