A rotary hydraulic actuator may be configured to output rotary motion to control a hinged surface of an aircraft. The actuator includes a nested ballscrew, ballnut, and output assembly that form concentric ball races for converting the linear motion and force of the linear actuator to rotary motion and torque of the output assembly that is connected to the hinged surface. One of the ball races is helically inclined and the other of the ball races is linear. The rotary hydraulic actuator may include a ball return structure that returns the balls from a loaded path of a ball race to an unloaded path of the ball race. The ball return structure may define a ball return path that is located at the same radial distance from the actuator centerline as the loaded path for minimizing the overall diameter of the actuator.

A rotary hydraulic actuator may be configured to output rotary motion to control a hinged surface of an aircraft. The actuator includes a nested ballscrew, ballnut, and output assembly that form concentric ball races for converting the linear motion and force of the linear actuator to rotary motion and torque of the output assembly that is connected to the hinged surface. One of the ball races is helically inclined and the other of the ball races is linear. The rotary hydraulic actuator may include a ball return structure that returns the balls from a loaded path of a ball race to an unloaded path of the ball race. The ball return structure may define a ball return path that is located at the same radial distance from the actuator centerline as the loaded path for minimizing the overall diameter of the actuator.

An automatic engine driven pump (EDP) depressurization system for an aircraft is disclosed. The aircraft includes at least two EDPs driven by a main engine for converting mechanical power provided by the main engine into hydraulic power for distribution by a hydraulic system. The EDP depressurization system includes a depressurization device corresponding to each of the at least two EDPs and a control module. The depressurization devices are each energized to depressurize a respective EDP. The control module is in signal communication with each of the depressurization devices. The control module includes control logic for automatically generating a depressurization signal that energizes one of the depressurization devices based on a plurality of operational conditions of the aircraft.

An automatic engine driven pump (EDP) depressurization system for an aircraft is disclosed. The aircraft includes at least two EDPs driven by a main engine for converting mechanical power provided by the main engine into hydraulic power for distribution by a hydraulic system. The EDP depressurization system includes a depressurization device corresponding to each of the at least two EDPs and a control module. The depressurization devices are each energized to depressurize a respective EDP. The control module is in signal communication with each of the depressurization devices. The control module includes control logic for automatically generating a depressurization signal that energizes one of the depressurization devices based on a plurality of operational conditions of the aircraft.

The present invention relates to an aircraft having at least one flap arranged at the wing of the aircraft and having at least one first drive unit for actuating the flap as a landing flap and a first control unit for controlling the first drive unit when the aircraft is in a landing mode of operation, wherein the aircraft comprises at least one second drive unit which is an active differential gear box for actuating the flap as an aileron and a second control unit for controlling the second drive unit when the aircraft performs an aileron function.

The present invention relates to an aircraft having at least one flap arranged at the wing of the aircraft and having at least one first drive unit for actuating the flap as a landing flap and a first control unit for controlling the first drive unit when the aircraft is in a landing mode of operation, wherein the aircraft comprises at least one second drive unit which is an active differential gear box for actuating the flap as an aileron and a second control unit for controlling the second drive unit when the aircraft performs an aileron function.

Methods, apparatus, systems and articles of manufacture are disclosed for redundant actuation of control surfaces. An example apparatus includes a control surface of an aircraft, and an actuator to move the control surface. The example apparatus also includes an electric motor to move the actuator; the electric motor communicatively coupled to an electrical system of the aircraft. The example apparatus also includes a hydraulic motor to move the actuator, the hydraulic motor fluidly coupled to a hydraulic system of the aircraft. The example apparatus also includes a sensor to detect incorrect operation of the hydraulic system. The example apparatus also includes a switch operatively coupled to the sensor, the switch to enable operation of the electric motor in response to the detected incorrect operation of the hydraulic system.

Methods, apparatus, systems and articles of manufacture are disclosed for redundant actuation of control surfaces. An example apparatus includes a control surface of an aircraft, and an actuator to move the control surface. The example apparatus also includes an electric motor to move the actuator; the electric motor communicatively coupled to an electrical system of the aircraft. The example apparatus also includes a hydraulic motor to move the actuator, the hydraulic motor fluidly coupled to a hydraulic system of the aircraft. The example apparatus also includes a sensor to detect incorrect operation of the hydraulic system. The example apparatus also includes a switch operatively coupled to the sensor, the switch to enable operation of the electric motor in response to the detected incorrect operation of the hydraulic system.

The present disclosure relates to a mechanical actuator having a modified locking mechanism and fewer components. The actuator has a cylinder, a locking recess formed on an interior wall of the cylinder, and a piston assembly that moves between an extended position and a retracted position responsive to fluid pressure within the cylinder. A lock is connected to the piston assembly and moves radially between a locked position and an unlocked position responsive to the fluid pressure within the cylinder.

An aircraft hydraulic actuation system for retracting an aircraft landing gear. The actuation system includes a supply line arranged to carry hydraulic fluid pressurized by a pump, a return line arranged to return hydraulic fluid to a reservoir, and a hydraulic actuator 128. In a first mode of operation, a first chamber 130 of the actuator 128 is supplied with pressurized hydraulic fluid from the supply line such that a piston 134 is moved in a first direction so as to move a load such as a landing gear. In a second mode of operation, the first chamber 130 is taken out of fluid communication with the supply line and a second chamber 132 is in fluid communication with the return line, such that the piston 134 is able to be moved under the influence of the load, for example when the landing gear extends under gravity.