Actuator in a landing gear system of an aircraft

1. Actuator in a landing gear system of an aircraft, comprising:

an electric drive for driving the actuator; and

first drive electronics for controlling the electric drive directly connected to the drive via an electric lined; and

second drive electronics for controlling the electric drive directly connected to the drive via an electric line, wherein

the first drive electronics and the second drive electronics are redundant with respect to one another and independently operate from one another,

the electric drive is an electric motor that is connected to both the first drive electronics and the second drive electronics,

the first drive electronics and the second drive electronics are directly connected to one another only via a communications link, and

the electric motor is a dual winding motor whose windings are electrically independent of one another, with one of the windings cooperating with the first drive electronics and the other winding cooperating with the second drive electronics,

and further comprising a switchover unit that permits a control of the electric motor by either the first drive electronics or the second drive electronics, the switchover unit being configured to be switched on upon failure of one of the first or second drive electronics,

such that the switchover unit disables communication from the failed drive electronics to the electric motor, and failure of one of the first or second drive electronics is compensated by the other of the first or second drive electronics.

2. An actuator in accordance with claim 1 , wherein the actuator is an electromechanical actuator or an electrohydraulic actuator; and the electric drive is a pump of a hydraulic circuit with an electrohydraulic actuator.

3. An actuator in accordance with claim 1 , wherein the first drive electronics are different from or identical to the second drive electronics.

4. An actuator in accordance with claim 1 , wherein the actuator only has the one electric drive and/or the actuator has a decentralized hydraulic circuit.

5. An actuator in accordance with claim 1 , wherein the dual winding is present on a common shaft and/or on the same rotor magnets.

6. An actuator in accordance with claim 1 , wherein the electric motor has redundantly implemented phases, with one of the redundant phases cooperating with the first drive electronics and the other one of the redundant phases cooperating with the second drive electronics.

7. An actuator in accordance with claim 1 , wherein the electric motor is a 3-phase permanent magnetic synchronous motor.

8. An actuator in accordance with claim 1 , wherein the actuator is adapted to actuate landing gear or to control a landing gear wheel.

9. An actuator in accordance with claim 1 , wherein the performance of the actuator or of the electric drive is independent of the failure of one of the two mutually redundant drive electronics.

10. An actuator in accordance with claim 1 , wherein the performance of the actuator or of the electric drive reduces on the failure of one of the two mutually redundant drive electronics.

11. A landing gear system of an aircraft that comprises a plurality of actuators in accordance with claim 1 , wherein the first drive electronics and the second drive electronics for the plurality of actuators of the landing gear system are arranged together in a decentralized manner at a location.

12. An actuator in accordance with claim 1 , wherein said actuator comprises only a single electric drive, in turn having two separate, redundant drive electronics.

13. An actuator in accordance with claim 12 , wherein only the electronics of the actuator are provided as redundant.

14. An actuator in accordance with claim 1 , wherein only the electronics of the actuator are provided as redundant.

15. An actuator in accordance with claim 7 , wherein the motor is provided with a resolver or with a Hall effect sensor for the motor regulation.

16. An actuator in accordance with claim 10 , wherein the failure of one of the two mutually redundant drive electronics is reduced by half.

17. An actuator in a landing gear system of an aircraft, comprising:

an electric drive for driving the actuator;

first drive electronics for controlling the electric drive directly connected to the drive via an electric line; and

second drive electronics for controlling the electric drive directly connected to the drive via an electric line, wherein

the first drive electronics and the second drive electronics are redundant with respect to one another and independently operate from one another,

the electric drive is an electric motor that is connected to both the first drive electronics and to the second drive electronics,

the first drive electronics and the second drive electronics are directly connected to one another only via a communications link, and

the electric motor is a dual winding motor whose windings are electrically independent of one another, with one of the windings cooperating with the first drive electronics and the other winding cooperating with the second drive electronics,

and further comprising a switchover unit that permits a control of the electric motor by either the first drive electronics or the second drive electronics, the switchover unit being configured to be switched on upon failure of one of the first or second drive electronics,

such that the switchover unit disables communication from the failed drive electronics to the electric motor, and failure of one of the first or second drive electronics is compensated by the other of the first or second drive electronics, and

a first motor sensor for determining an operating state of the drive and a second motor sensor for determining an operating state of the drive, with the first motor sensor being electrically connected to the first drive electronics and the second motor sensor being electrically connected to the second drive electronics.