A gas turbine engine according to an example of the present disclosure includes, among other things, a fan including a plurality of fan blades rotatable about an engine axis. Each of the fan blades have a leading edge and rotate about a fan blade axis. A method of operating a gas turbine engine is also disclosed.

A fan assembly for a heating, ventilation, and/or air conditioning (HVAC) unit includes a hub having a fan blade extending therefrom. The fan assembly includes a fan pitch adjuster configured to engage the fan blade and to adjust a pitch of the fan blade relative to the hub in response to movement of the fan pitch adjuster relative to the hub. The fan assembly also includes a motor shaft coupled to the fan pitch adjuster and to the hub, where the motor shaft is configured to drive rotation of the fan pitch adjuster and the hub. The fan assembly further includes an actuator coupled to the fan pitch adjuster and configured to axially move the fan pitch adjuster along the motor shaft and relative to the hub to adjust the pitch of the fan blade.

There is provided a blade angle feedback system for an aircraft-bladed rotor rotatable about a longitudinal axis and having an adjustable blade pitch angle. A feedback device is coupled to rotate with the rotor and to move along the axis with adjustment of the blade pitch angle. The feedback device comprises a body having position marker(s) embedded therein, the body made of a first material having a first magnetic permeability and the position marker(s) comprising a second material having a second magnetic permeability greater than the first. Sensor(s) are positioned adjacent the feedback device and configured for producing, as the feedback device rotates about the axis, sensor signal(s) in response to detecting passage of the position marker(s). A control unit is communicatively coupled to the sensor(s) and configured to generate a feedback signal indicative of the blade pitch angle in response to the sensor signal(s) received from the sensor(s).

A blade retainer adapted for use in a gas turbine engine is configured to block axial movement of a blade. The blade retainer includes a first brace, a second brace, and a web that extends between the first brace and the second brace. The blade retainer is configured to block axial movement of a root of the blade out of a blade receiver slot.

Gas turbine aircraft engine comprising an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor; and a fan upstream of the engine core and driven by the turbine, the fan comprising a circumferential row of tandem fan blades. Each fan blade comprises a main blade and an auxiliary blade. Over substantially all of the auxiliary blade's radial span, the leading edge of the auxiliary blade is rearwards of the closest point on the trailing edge of the main fan blade, and on a given aerofoil chordal section of the main fan blade, the leading edge position of an aerofoil chordal section of the auxiliary fan blade lies on a rearwards extension of the camber line of the aerofoil chordal section of the main fan blade, and the main fan blade and the auxiliary fan blade are arranged to rotate in tandem.

A bearing system is configured to surround a rotor along a circumferential direction corresponding to the rotor. The rotor is extended along an axial direction co-directional to a centerline axis of the rotor. The bearing system includes a body from which a plurality of first support members is each extended, and wherein the plurality of first support members is spaced apart from one another along the circumferential direction. Each first support member includes a first axial support arm extended along the axial direction and a first radial support arm extended from the first axial support arm along a radial direction. The first radial support arm is configured to position a bearing element in contact with a bearing surface at the rotor.

There are described methods and systems for operating an aircraft turboprop engine. The method comprises controlling a propeller of the turboprop engine based on a selected one of a reference propeller rotational speed and a minimum propeller blade angle while the turboprop engine is running; detecting an inflight restart of the turboprop engine; and controlling the propeller during the inflight restart in accordance with at least one of a modified reference propeller rotational speed and a modified minimum propeller blade angle to maintain an actual propeller blade angle above an aerodynamic disking angle during the inflight restart.

A propeller assembly of an aircraft includes a hub, a plurality of propeller blades extending from the hub and secured thereto and a propeller blade pitch change system located at at least one propeller blade of the plurality of propeller blades. The propeller blade pitch change system includes a pitch change actuator located in the propeller blade, and a drive mechanism operably connected to the pitch change actuator and to the propeller blade to urge rotation of the propeller blade about a propeller blade axis.

Embodiments of a direct-drive fan system and a variable process control system are disclosed herein. The direct-drive fan system and the variable process control system efficiently manage the operation of fans in a cooling system such as a wet-cooling tower or air-cooled heat exchanger (ACHE), HVAC systems, mechanical towers or chiller systems.

An aspect includes a variable cycle system of a gas turbine engine. The variable cycle system includes an actuation system, an electric component, and a controller. The actuation system is configured to adjust a variable cycle of turbomachinery of the gas turbine engine. The electric component is operable to provide a shaft power supply or a load corresponding respectively to an adjustment of the turbomachinery. The controller is operable to adjust an output of either or both of the actuation system and the electric component for separate control of thrust and cycle responses.