A gas turbine engine includes a high speed turbine and a high speed compressor connected by a high speed shaft to define a high speed spool. A low speed turbine and a low speed compressor are connected by a low speed shaft to define a low speed spool. A high speed tower shaft is driven by the high speed spool and a low speed tower shaft is driven by the low speed spool. The low speed tower shaft drives a low speed input gear in an accessory drive differential. The high speed tower shaft drives a high speed input gear in the accessory drive differential. One of the high speed input gear and the low speed input gear increases an output speed of an output gear, and the other decreases an output speed of the output gear. The output gear drives at least one accessory for the gas turbine engine.

A gas turbine engine includes a high speed turbine and a high speed compressor connected by a high speed shaft to define a high speed spool. A low speed turbine and a low speed compressor are connected by a low speed shaft to define a low speed spool. A high speed tower shaft is driven by the high speed spool and a low speed tower shaft is driven by the low speed spool. The low speed tower shaft drives a low speed input gear in an accessory drive differential. The high speed tower shaft drives a high speed input gear in the accessory drive differential. One of the high speed input gear and the low speed input gear increases an output speed of an output gear, and the other decreases an output speed of the output gear. The output gear drives at least one accessory for the gas turbine engine.

An aspect includes a hybrid gas turbine engine system of a hybrid electric aircraft. The hybrid gas turbine engine system includes a gas turbine engine having a low speed spool and a high speed spool, a generator operably coupled to the low speed spool, a high spool electric motor operably coupled to the high speed spool, and a controller. The controller is configured to control the hybrid gas turbine engine system in a powered warm-up state to add heat to one or more components of the gas turbine engine by operating the gas turbine engine with a higher engine power setting above idle to drive rotation of the generator, transfer power from the generator to the high spool electric motor, and produce thrust. The gas turbine engine transitions from the powered warm-up state after reaching a target temperature of the one or more components in the powered warm-up state.

An aspect includes a hybrid gas turbine engine system of a hybrid electric aircraft. The hybrid gas turbine engine system includes a gas turbine engine having a low speed spool and a high speed spool, a generator operably coupled to the low speed spool, a high spool electric motor operably coupled to the high speed spool, and a controller. The controller is configured to control the hybrid gas turbine engine system in a powered warm-up state to add heat to one or more components of the gas turbine engine by operating the gas turbine engine with a higher engine power setting above idle to drive rotation of the generator, transfer power from the generator to the high spool electric motor, and produce thrust. The gas turbine engine transitions from the powered warm-up state after reaching a target temperature of the one or more components in the powered warm-up state.

A gas turbine propulsion system includes a shroud that defines a fluid flow path. A gas turbine engine in the fluid flow path includes a compressor, a combustor downstream from the compressor, and a turbine downstream from the combustor. An electric generator in the fluid flow path includes a rotor coaxially aligned with the turbine. A propulsor is upstream from the gas turbine engine, and an electric motor is operably coupled to the propulsor to rotate the propulsor. The propulsor is rotationally isolated from the gas turbine engine so that the propulsor rotates independently from operation of the gas turbine engine.

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 aircraft gas turbine engine comprises at least a first main engine shaft an electric machine coupled to the main engine shaft, and a lubrication pump configured to provide oil to at least a part of the electric machine. The lubrication pump is controlled by a controller. The controller is configured to operate the lubrication pump in a main gas turbine engine running mode, in which the lubrication pump is driven to provide lubrication to the electric machine at a first flow rate relative to the rotational rate of the main engine shaft, and a fault mode, in which the lubrication pump is driven to provide lubrication to the electric machine at a second, higher flow rate relative to the rotational rate of the main engine shaft.

A turbomachine has an electric machine, and a starter-generator incorporating a second electric machine configured to operate as a starter and a generator. The the electric machine and the generator-starter are arranged to transfer electric power between there. A planetary gear reduction device has a first element coupled to an accessory relay housing, a second element coupled to the starter-generator, a third element rotated by the electric machine, control means configured to modify the speed of rotation of the first electric machine and the starter-generator so as to provide the starter-generator with a maximum torque in starter mode. The planetary gear reduction device also modifies the speed of rotation of the first electric machine such that the starter-generator is driven at a constant speed in generator mode.

A turbomachine has an electric machine, and a starter-generator incorporating a second electric machine configured to operate as a starter and a generator. The the electric machine and the generator-starter are arranged to transfer electric power between there. A planetary gear reduction device has a first element coupled to an accessory relay housing, a second element coupled to the starter-generator, a third element rotated by the electric machine, control means configured to modify the speed of rotation of the first electric machine and the starter-generator so as to provide the starter-generator with a maximum torque in starter mode. The planetary gear reduction device also modifies the speed of rotation of the first electric machine such that the starter-generator is driven at a constant speed in generator mode.