Methods and systems for starting an aircraft propulsion engine are described herein. Operation of the engine is controlled by an engine controller having a first channel and a second channel. An engine start request is received on the second channel while the first channel is inactive. It is determined which one of the first channel and the second channel was used to conduct a last successful engine start. Responsive to determining that the last successful engine start was conducted on the second channel, a predetermined time period is waited to elapse and for the first channel to reinitialize before commanding, via the first channel, an engine start based on the engine start request.

A blade tip clearance control apparatus is disclosed. The apparatus includes a rotor, a hydraulic clearance control device, and a cylinder locking device. The rotor includes a thrust collar, a pair of thrust bearings axially supporting the thrust collar, and a plurality of radially extending blades. The hydraulic clearance control device includes a first hydraulic cylinder moving any one of the pair of thrust bearings in a forward axial direction and a second hydraulic cylinder moving the other thrust bearing in the reverse axial direction. The cylinder locking device includes a first locking device restricting a forward moving distance of the first hydraulic cylinder and a second locking device restricting a reverse moving distance of the second hydraulic cylinder.

A gas turbine engine with a geared turbofan arrangement with a gearbox in a drive train driven by a turbine, a driving side of the gearbox being driveably connected with a propulsive fan, is provided. The gas turbine includes at least one form locking connection device in a drive train enabling a controlled disengagement of at least one engine part from the drive train in case of a mechanical failure of the gas turbine engine or a part thereof and wherein the at least one form locking connection device is positioned in a torque carrying shaft or a torque carrying part of a shaft and/or wherein the at least one form locking connection device is positioned between the torque bearing coupling of the gearbox with the fan shaft and a torque carrier of the gearbox and at least one load stop for bearing an essential axial load.

A control assembly for an aircraft system according to an example of the present disclosure includes a multi-core processor that has a plurality of cores coupled to a communications module and to an arbitration module. The communications module is operable to communicate information between the plurality of cores and one or more aircraft modules. The plurality of cores include first and second cores operable to concurrently execute a first discrete set of software instructions to generate respective instances of an output. The arbitration module is operable to communicate each and every one of the respective instances to control the one or more aircraft modules. A method of operating an aircraft system is also disclosed.

An engine control system includes an electronic hardware engine controller and an actuator that operates at different positions to control operation of an engine. An actuator sensor measures an actuator position, and the engine controller generates a synthesized actuator position. In response to detecting a faulty actuator, a faulty actuator sensor, or both, the engine controller adjusts the position of the actuator based on the synthesized actuator position.

Systems and methods include actuation system of turbomachinery that include an actuator configured to control pitch of vanes of the turbomachinery and a positioner configured to position the actuator to control the pitch. The actuation system also includes a controller system communicably coupled to the positioner. The controller system is configured to drive the positioner based at least in part on a set point of the turbomachinery and obtain data from the positioner indicating an amount of force used for a target pitch. Using the data, the controlling system derives an operating condition for the turbomachinery and controls the turbomachinery based at least in part on the operating condition of the turbomachinery.

A method includes controlling an electric motor of a hybrid-electric powerplant for an aircraft using an EPC (electric powertrain controller) and controlling a heat engine of the hybrid-electric powerplant using an ECU (engine control unit). The method includes performing at least one of the following to protect the hybrid-electric powerplant: using the ECU to power down the electric motor, and/or using the EPC to power down the heat engine.

A lubrication system for a gas turbine engine, the system comprising: a gearbox, the gearbox comprising a sump; an oil tank; a primary gearbox lubrication system configured to pump oil from the oil tank to lubricate the gearbox with a gearbox primary feed; a secondary gearbox lubrication system, configured to lubricate the gearbox with oil from the sump when the oil level in the sump reaches a predetermined level; wherein the system is configured to increase the oil level in the sump to at least the predetermined level in response to a failure of the primary gearbox lubrication system.

A variable pitch fan for gas turbine engine is provided. The variable pitch fan includes a plurality of fan blades rotatably coupled to a disk and an actuation assembly for changing a pitch of each of the plurality of fan blades. The actuation assembly generally includes an actuation member operably connected to at least one of the plurality of fan blades and a pin. The pin is movable between a first position in which the pin is positioned at least partially in a channel defined in the actuation member, and a second position. The pin blocks movement of the actuation member relative to the pin past a range defined by the channel when in the first position. The actuation assembly further includes a retraction system for selectively engaging the pin in moving the pin from the first position to the second position.

An actuation system is provided for varying the pitch of the blades of a variable pitch fan or propeller. The actuation system includes a first actuator and a second actuator. The actuation system further includes a first linkage which operably connects the first actuator to the blades such that operation of the first actuator varies the pitch of the blades via the first linkage. The actuation system further includes a second linkage which operably connects the second actuator to the first actuator such that operation of the second actuator varies the pitch of the blades by series connection of the second linkage to the first linkage.