An accessory drive for an engine includes a power takeoff (PTO) configured to couple power from a rotating shaft of the engine and to convey the power through an opening in a housing of the engine. A gearbox is coupled to and configured to be driven by the PTO. The gearbox is disposed external to the housing and includes a planetary gear train. At least one engine accessory is coupled to and configured to be driven by the planetary gear train.

An APU has a gas turbine engine and a starter generator to be selectively driven by the gas turbine engine. A sensor senses windmilling of components associated with the starter generator. A lock feature limits rotation within the starter generator when windmilling is sensed. A method of operation is also disclosed.

An APU has a gas turbine engine and a starter generator to be selectively driven by the gas turbine engine. A sensor senses windmilling of components associated with the starter generator. A lock feature limits rotation within the starter generator when windmilling is sensed. A method of operation is also disclosed.

The invention relates to a removable reactivation pack for a turboshaft engine of a helicopter, comprising a gas generator equipped with a drive shaft, said turboshaft engine (6) being capable of operating in at least one standby mode during a stable flight of the helicopter, said removable pack comprising: a removable gearbox comprising a gearbox output shaft; controlled means for rotating said gearbox output shaft, referred to as reactivation means of said turboshaft engine; mechanical means for reversibly coupling said gearbox output shaft to said drive shaft of said gas generator.

A hybrid-electric propulsion system for an aircraft is provided herein that can include a propulsor and a turbomachine comprising a high pressure turbine drivingly coupled to a high pressure compressor through a high pressure spool. An auxiliary power unit can be operably coupled with a starter motor. An electrical system can comprise a first electric machine coupled to the turbomachine. The first electric machine can be separate from the starter motor. A controller can be configured to provide electrical power from an electric power source to the first electric machine to drive the first electric machine to start, or assist with starting, the turbomachine.

A hybrid-electric propulsion system for an aircraft is provided herein that can include a propulsor and a turbomachine comprising a high pressure turbine drivingly coupled to a high pressure compressor through a high pressure spool. An auxiliary power unit can be operably coupled with a starter motor. An electrical system can comprise a first electric machine coupled to the turbomachine. The first electric machine can be separate from the starter motor. A controller can be configured to provide electrical power from an electric power source to the first electric machine to drive the first electric machine to start, or assist with starting, the turbomachine.

A system and a method for the automatic starting of the engines of a twin-engine aircraft. When a first engine is made by the pilots of the aircraft to follow an automatic starting procedure, this engine follows the conventional four-phase starting cycle. The rotor of the high-pressure stage of the second engine is spun up for a predetermined duration and to a speed that is substantially constant and less than the nominal value during the starting cycle for the first engine. Once the starting cycle for the first engine is complete, the second engine is made by the pilots of the aircraft to follow an automatic starting procedure. The starting cycle for the second engine comprises just three phases.

A bowed rotor start mitigation system for a gas turbine engine is provided. The bowed rotor start mitigation system includes a controller operable to receive a vibration input indicative of an actual vibration level of the gas turbine engine and a speed input indicative of a rotor speed of the gas turbine engine. The controller generates a bowed rotor start mitigation request based on comparing the actual vibration level to a modeled vibration level at the rotor speed.

A bowed rotor start mitigation system for a gas turbine engine is provided. The bowed rotor start mitigation system includes a controller operable to receive a vibration input indicative of an actual vibration level of the gas turbine engine and a speed input indicative of a rotor speed of the gas turbine engine. The controller generates a bowed rotor start mitigation request based on comparing the actual vibration level to a modeled vibration level at the rotor speed.

A bowed rotor start mitigation system for a gas turbine engine of an aircraft is provided. The bowed rotor start mitigation system includes a motoring system and a controller coupled to the motoring system and an aircraft communication bus. The controller is configured to determine at least one inferred engine operating thermal parameter from at least one aircraft-based parameter received on the aircraft communication bus. The motoring system is controlled to drive rotation of a starting spool of the gas turbine engine below an engine idle speed based on determining that the at least one inferred engine operating thermal parameter is within a preselected threshold.