In a rotating machine having a fluidic rotor, the rotor comprises at least one blade mounted on an arm rotating about a rotor shaft forming a main axis of the rotor, the rotor being kept by a supporting structure in an orientation such that said axis is substantially perpendicular to the direction of flow of the fluid, the blade being mounted so as to pivot about an axis of rotation of the blade parallel to the main axis. The machine comprises means for generating a relative oscillation movement of the blade with respect to the arm at the axis of rotation of the blade, in order in this way to vary the inclination of the blade during the rotation of the rotor. Said means comprise, at the arm end, a mechanism comprising a first rotating element (A; B) known as the drive element and a second rotating element (B; A) known as the driven element, the elements being mounted on mutually parallel axes of rotation and separated by an inter-axis distance, the orientation of the drive element being controlled depending on the orientation of the rotor shaft while the orientation of the driven element determines the orientation of the blade, one of the rotating elements comprising a finger (D) spaced apart from its axis of rotation and the other rotating element comprising a groove (C) which receives the finger and in which the finger can slide. Application notably to wind turbines, to marine turbines and to nautical and aircraft propellers.

Apparatus for actuating and controlling the rotation, about their longitudinal axis (Y-Y), of blades (20) of cooling fans (10) for operating machines and/or vehicles, in particular agricultural tractors and off-road vehicles, said fan being mounted on a hub (11) which can be rotationally driven about its axis (X-X) by associated driving means (3,3a) suitable for connection to the heat engine (1) and mounted on a fixed support (5) by means of a bearing (3b) the apparatus comprising a ring (71) provided with a radial seat (71a) inside which a radial pin (72), eccentrically engaged in a base (73) integral with the shank (20a) of the blade (20), is inserted; an electric motor (30) which is coaxial with the axis (X-X) of the hub (11) and the shaft (31) of which is coaxially connected to a reduction gear (40), the kinematic output element (143) of which is coaxially connected by means of a screw (51a)/female thread (76a) coupling to a slider (76) displaceable in both directions along the axis (X-X) and kinematically connected to the ring (71) with an eccentric pin (72) driving the base (73) of the shank of the blade, an electromagnetic clutch (80; 180; 280) being arranged between the pulley (3) and the hub (11) of the fan.

An actuator is disclosed comprising: a moveable member; a first hydraulic chamber in contact with a first surface of the moveable member; a second hydraulic chamber in contact with a second, opposing surface of the moveable member; a movable locking mechanism coupled to a moveable wall of the second hydraulic chamber; and a resilient biasing member acting on the moveable wall of the second hydraulic chamber so as to bias the moveable wall and locking mechanism. The actuator is configured to selectively vary the pressure in the second hydraulic chamber so that the resilient biasing member is able to bias the moveable wall to move, thereby moving the locking mechanism to engage the moveable member so as to prevent the movement of the moveable member towards at least one of the first and second hydraulic chambers.

A propeller oil control system for a turboprop engine of an aircraft includes an engine control unit and a propeller oil controller. The engine control unit is operable to determine a flight phase of the aircraft and is configured to supply control commands. The propeller oil controller is coupled to receive a supply of oil and to discharge the oil at a discharge oil pressure. The propeller oil controller includes an electrohydraulic servo valve that receives the control commands moves to a plurality of positions between a first position and a mid-position, and a plurality of positions between the mid-position and a second position. The engine control unit only commands the electrohydraulic servo valve to move out of the first position when the engine control unit determines the aircraft is conducting a take-off roll or the aircraft is in flight.

A system for controlling a propeller having a plurality of blades having a primary control system and a backup control system. The primary control system including a sensor responsive to a propeller state, and a controller connected to the sensor and to an electrohydraulic control actuator. The electrohydraulic control actuator connected via a bypass valve to a hydraulic actuator that controls at least a blade angle of a blade of the propeller. The controller generating commands to the electrohydraulic control actuator based on at least the propeller state. The backup control system including a second controller, an electrohydraulic solenoid operably connected to the bypass valve. The backup control system operable to hydraulically disable the primary control system via the bypass valve upon the occurrence of a selected condition, the second controller modulates the operation of the electrohydraulic solenoid to control the bypass actuator based on the propeller state.

Hydraulic pitch actuation mechanism includes vaned rotor within timing chamber surrounded by annular timing chamber wall within fan hub and variable area and volume retarding and advancing chambers within timing chamber. Timing pocket walls extend inwardly from timing chamber wall and interdigitated with timing vanes extending outwardly from vane shaft of vaned rotor. Hydraulic retarding and advancing fluid passages extend through fan drive shaft and through the fan hub to the retarding and advancing chambers respectively. Fluid passages include annular axially spaced apart retarding and advancing passage discharge sections, annular and axially spaced apart retarding and advancing passage entry sections spaced apart from and aft of retarding and advancing passage discharge sections respectively. Retarding and advancing connecting passage sections extend through fan drive shaft and fluidly connect retarding and advancing passage entry sections to retarding and advancing passage discharge sections. Hub passages through fan hub connect passage discharge sections to chambers.

Energy scavenging health monitors are provided for assessing the health of components onboard aircraft and other vehicles, as are methods carried-out by energy scavenging health monitors. In various embodiments, the energy scavenging health monitor includes an energy scavenger system, a controller coupled to the energy scavenger system, and a first sensor coupled to the controller. During operation of the health monitor, the first sensor provides sensor signals to the controller, which are indicative of an operational parameter pertaining to a monitored device of the vehicle. Storage media contains computer-readable instructions, which when executed by the controller, cause the energy scavenging health monitor to determine when a predetermined trigger event has occurred based, at least in part, on electrical input signals received from the energy scavenger system. In response to occurrence of the predetermined trigger event, the controller further stores data indicative of the sensor signals in the storage media.

A gas turbine engine including a core having in serial flow order a compressor, a combustor, and a turbinethe compressor, combustor, and turbine together defining a core air flowpath. The gas turbine engine additionally includes a fan section mechanically coupled to the core, the fan section including a plurality of fan blades, and each of the plurality fan blades defining a pitch axis. An actuation device is operable with the plurality fan blades for rotating the plurality fan blades about their respective pitch axes, the actuation device including an actuator located outward of the core air flowpath to, e.g., simplify the gas turbine engine.

A method and system for controlling a pitch of blades of a fan assembly having a centerline axis of rotation is provided. The system includes a pitch change mechanism (PCM) including a hydraulic actuator positioned axisymmetric with respect to the fan assembly and configured to angularly displace the blades of the fan assembly between a first position and a second position. The PCM further includes a plurality of hydraulic fluid supply lines coupled in flow communication between the hydraulic actuator and a hydraulic fluid transfer sleeve, the hydraulic fluid transfer sleeve configured to transfer a flow of pressurized hydraulic fluid across a gap between a stationary member of the hydraulic fluid transfer sleeve and a rotatable member of the hydraulic fluid transfer sleeve.

A system and method for determining a position of a feedback ring of a propeller of an aircraft engine are provided. The feedback ring is coupled to the propeller to rotate with the propeller and to be displaced along a longitudinal axis with adjustment of a blade angle. An engagement member is configured to engage the feedback ring and to be displaced along a longitudinal direction substantially parallel to the longitudinal axis with displacement of the feedback ring. A sensor comprises a first member coupled to the engine and a second member coupled to the engagement member. The second member is moveable relative to the first member along the longitudinal direction as the engagement member is displaced. The sensor generates a signal indicative of a longitudinal position of the second member relative to the first member. A controller determines an axial position of the feedback ring from the sensor signal.