The force feedback device includes a stationary housing intended to be secured to an understructure of an aircraft, at least one spring opposing movement of a side-stick of the aircraft relative to the housing, and a transmission mechanism, which is supported by the housing movably and which is suitable for transmitting a rotational movement, around a first rotation axis, between the side-stick and said at least one spring, by applying a force law according to which a resistive force, opposed by said at least one spring, via the transmission mechanism, against the rotational movement of the side-stick around the first rotation axis, depends on an angular position of the side-stick around the first rotation axis. In order for this force feedback device to be more precise, compact and reliable, the transmission mechanism comprises a cam, which is suitable for being connected to the side-stick in rotation around the first rotation axis and which is provided with at least one profiled surface that is shaped so as to define at least one part of the force law, said at least one part of the force law including at least one force jump that corresponds to a discontinuity in intensity of the resistive force without changing the direction of the latter.

A control system (200) using fixed point computation avoids overflow conditions by limiting the internal quantities ensuring that contributing calculations from various parts of the control loop never add up to an overflow. Various parts of a control loop e.g. proportional terms can be prioritised over other parts, e.g. integral and differential terms and ensuring that the overall total resulting when the separate terms are summed together never exceeds the maximum or minimum imposed limits. Variable limit calculator circuits (211,219) revise the limits according to the output of the higher priority control path(s). The revised limits cascade down through each contributor, eventually allowing the lowest priority control path the smallest authority. The control system may be applied to control of a drive motor (111) for an aircraft inceptor (103).

The invention provides a device comprising a control lever associated with a body carrying a first plate connected to the body, a second plate connected to the first plate with the control lever being connected to the second plate, a first transmission shaft pivotally mounted relative to the body, a first connection mechanism connecting the control lever to the first transmission shaft in such a manner that pivoting of the control lever about the first axis causes the first shaft to pivot about the fourth axis, a second transmission shaft pivotally mounted relative to the body, a second connection mechanism connecting the control lever to the second transmission shaft in such a manner that pivoting of the control lever about the second axis causes the second shaft to pivot about the sixth axis, and a platform connected to the body, the second connection mechanism being pivotally mounted on the platform and being pivotally mounted on the control lever.

A helicopter autopilot system includes an inner loop for attitude hold for the flight of the helicopter including a given level of redundancy applied to the inner loop. An outer loop is configured for providing a navigation function with respect to the flight of the helicopter including a different level of redundancy than the inner loop. An actuator provides a braking force on a linkage that serves to stabilize the flight of the helicopter during a power failure. The actuator is electromechanical and receives electrical drive signals to provide automatic flight control of the helicopter without requiring a hydraulic assistance system in the helicopter. The autopilot can operate the helicopter in a failed mode of the hydraulic assistance system. A number of flight modes are described with associated sensor inputs including rate based and true attitude modes.

A control stick pivot (200) comprises a primary pivot bearing (210), a redundant pivot bearing (220), and a detection component for detecting malfunction of the primary pivot bearing. The redundant pivot bearing is provided on one side of the primary pivot bearing. The redundant pivot bearing becomes operative as a bearing in the event that the primary pivot bearing malfunctions. The detection component may comprise one or more features (230) provided on the redundant pivot bearing for applying a torque to the redundant pivot bearing during maintenance or testing and/or a sensor (300; 310; 320) configured to detect relative motion at an interface between the primary pivot bearing and the redundant pivot bearing.

The present invention relates to an electronic gear shifter assembly for a dual-mode flying and driving vehicle. The electronic gear shifter assembly may include a lever moveable between a first shifting path that includes at least one drive-related operating position, and a second shifting path that includes at least one flying-related operating position.

Disclosed are systems and methods for controlling an electric vertical take-off and landing (eVTOL) aircraft. In one embodiment, a system comprises a processor, a first inceptor, communicatively coupled to the processor, the first inceptor configured to accept longitudinal and lateral linear movements as manual input and provide corresponding signals to the processor, and a second inceptor, communicatively coupled to the processor, the second inceptor configured to accept longitudinal and lateral linear movements as manual input and provide corresponding signals to the processor, wherein the processor is configured to control a heading of an aircraft using a signal received from the second inceptor corresponding to lateral linear movement of the second inceptor. Some embodiments may additionally include at least one sensor and a thumb stick for each inceptor.

Flight control systems, flight control laws, and aircraft are provided. An flight control system includes an input configured to receive a pitch rate command, a processor operative to receive the pitch angle command, to calculate a pitch angle saturation limit, to compare the sum of the pitch rate command, the scaled pitch rate, and the scaled pitch angle to the pitch angle saturation limit, to convert the pitch rate command system to the pitch angle command system in response to the sum exceeding the pitch angle saturation limit value to limit the pilot pitch-up pitch rate command, and to couple the pitch rate command to an aircraft control surface for the failure case of one of control surface, and the aircraft control surface configured to adjust an aircraft control surface setting in response to the pitch rate command and/or pitch angle command to protect an aircraft from being in stall condition.

Plinth system (1) for a control stick (51) for controlling an aircraft having a plinth flange (3) includes a tubular receiving means (7) and a plinth nut (5) which at least partially encloses the receiving means of the plinth flange (3), wherein the plinth nut (5) comprises an interface (11) for fastening a knurled nut (41).

The unified command system and/or method includes an input mechanism, a flight processor that receives input from the input mechanism and translates the input into control output, and effectors that are actuated according to the control output. The system can optionally include: one or more sensors, a vehicle navigation system which determines a vehicle state and/or flight regime based on data from the one or more sensors, and a vehicle guidance system which determines a flightpath for the aircraft.