A system for starting a gas turbine engine of an aircraft is provided. The system includes a pneumatic starter motor, a discrete starter valve switchable between an on-state and an off-state, and a controller operable to perform a starting sequence for the gas turbine engine. The starting sequence includes alternating on and off commands to an electromechanical device coupled to the discrete starter valve to achieve a partially open position of the discrete starter valve to control a flow from a starter air supply to the pneumatic starter motor to drive rotation of a starting spool of the gas turbine engine below an engine idle speed.

A bowed rotor prevention system for a gas turbine engine of an aircraft is provided. The bowed rotor prevention system includes a gear train and a bowed rotor prevention motor. The gear train is coupled through an engine accessory to an engine accessory gearbox that is further coupled to a starting spool of the engine. The bowed rotor prevention motor is operable to drive rotation of the starting spool of the gas turbine engine through the gear train upon engine shutdown.

A bowed rotor prevention system for a gas turbine engine of an aircraft is provided. The bowed rotor prevention system includes a gear train and a bowed rotor prevention motor. The gear train is coupled through an engine accessory to an engine accessory gearbox that is further coupled to a starting spool of the engine. The bowed rotor prevention motor is operable to drive rotation of the starting spool of the gas turbine engine through the gear train upon engine shutdown.

A method of operating an aircraft engine having a bladed rotor coupled thereto during an icing condition is provided. The method comprises controlling the aircraft engine to alternatingly achieve an ice accretion rotational speed of the bladed rotor at which ice accretion on blades of the bladed rotor occurs, and a reduced rotational speed of the bladed rotor to promote ice shedding from the blades of the bladed rotor. The reduced rotational speed is lower than the ice accretion rotational speed.

A method of operating an aircraft engine having a bladed rotor coupled thereto during an icing condition is provided. The method comprises controlling the aircraft engine to alternatingly achieve an ice accretion rotational speed of the bladed rotor at which ice accretion on blades of the bladed rotor occurs, and a reduced rotational speed of the bladed rotor to promote ice shedding from the blades of the bladed rotor. The reduced rotational speed is lower than the ice accretion rotational speed.

A sensing assembly comprising includes one or more sensing elements formed as a tube. The sensing assembly also includes one or more clamps to secure the one or more sensing elements. Each of the one or more clamps includes a recess in which each of the one or more sensing elements is seated.

A thrust reverser includes: a thrust-reversing element movable between a stowed position and a deployed position; at least one hydraulic actuator operably coupled to move the thrust-reversing element between the stowed and deployed positions; at least one hydraulic primary lock configured to transition, in response to a first activation pressure, between an engaged state, where movement of the thrust-reversing element is inhibited, and a released state, where movement of the thrust-reversing element is uninhibited; and a directional control unit fluidly coupled to the hydraulic actuator and the hydraulic primary lock, the directional control unit configured to transition from a first stage to a second stage in response to a second activation pressure that is greater than the first activation pressure, and where a transition from the first stage to the second stage by the directional control unit causes the hydraulic actuator to move the thrust-reversing element to the deployed position.

A thrust reverser includes: a thrust-reversing element movable between a stowed position and a deployed position; at least one hydraulic actuator operably coupled to move the thrust-reversing element between the stowed and deployed positions; at least one hydraulic primary lock configured to transition, in response to a first activation pressure, between an engaged state, where movement of the thrust-reversing element is inhibited, and a released state, where movement of the thrust-reversing element is uninhibited; and a directional control unit fluidly coupled to the hydraulic actuator and the hydraulic primary lock, the directional control unit configured to transition from a first stage to a second stage in response to a second activation pressure that is greater than the first activation pressure, and where a transition from the first stage to the second stage by the directional control unit causes the hydraulic actuator to move the thrust-reversing element to the deployed position.

A turbofan engine includes fan section including a plurality of fan blades, a gear train, a low spool including a low pressure turbine and a low pressure compressor, the low pressure turbine driving the plurality of fan blades through the gear train, and a high spool including a high pressure turbine driving a high pressure compressor. A fan nacelle at least partially surrounds a core nacelle to define a fan bypass flow path. A fan variable area nozzle is in communication with the fan bypass flow path and defines a fan nozzle exit area between the fan nacelle and the core nacelle. The fan variable area nozzle varies the fan nozzle exit area.

A turbofan engine includes fan section including a plurality of fan blades, a gear train, a low spool including a low pressure turbine and a low pressure compressor, the low pressure turbine driving the plurality of fan blades through the gear train, and a high spool including a high pressure turbine driving a high pressure compressor. A fan nacelle at least partially surrounds a core nacelle to define a fan bypass flow path. A fan variable area nozzle is in communication with the fan bypass flow path and defines a fan nozzle exit area between the fan nacelle and the core nacelle. The fan variable area nozzle varies the fan nozzle exit area.