KR101885663B1 - Fly-by-wire flight control system with backup mechanical flight control system - Google Patents

Abstract

The present invention relates to an electronic flight control system capable of electronically controlling a flying object in accordance with an operation unit command generated by a flight control computer and a flight control system having a mechanical control device capable of mechanically controlling the flying object as a backup control device . According to the present invention, even if a malfunction occurs in the electronic flight control system, the pilot can drive the rotor operating part by adjusting the pitch axis, the roll axis, the collective axis and the yaw axis of the air vehicle through the mechanical steering device, The safety of the pilots can be secured.

Description

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to an electronic flight control system,

The present invention relates to an electronic flight control system having a mechanical control device as a backup, and more particularly, to an electronic flight control system capable of electronically controlling an aircraft in accordance with an operation unit command generated by a flight control computer (FLY-BY-WIRE (FBW) SYSTEM) equipped with a mechanical control device capable of mechanically controlling a flying object as a backup control device.

Flight control systems are safety critical items that can affect the safety of aviation and pilots, requiring a very high level of safety.

In order to maximize flight safety, mission efficiency, and flight control performance, researches are actively conducted to apply electronic flight control systems to aircraft such as helicopters. However, rather than developing a completely new electronic flight control system, it would be better to retrofit a mechanical flight control system (MFCS) with an existing mechanical control system and limited limited authority into an electronic flight control system Cost and time savings, and reduced risk.

The electronic flight control system has the advantage that the control law can be designed so that it is convenient for the pilot to control the flight according to the flight characteristics of the flight control computer because all of the control authority for operating the rotor of the flight is controlled by the flight control computer. Accordingly, the electronic flight control system has been applied to military helicopters such as NH-90, UH-60M and CH-47, and civilian helicopters such as the S-92F. In addition, NRC's Bell 412 ASRA, NASA's RASCAL HACT, and DLT's ACT-FHS.

In the project to apply the electronic flight control system to the aircraft, the mechanical flight control system installed in the aircraft is removed, and the design is adopted by adopting only the electronic flight control system. However, in this case, if a malfunction occurs in the electronic flight control system, there is a problem that the pilots have very different means of controlling the aircraft, which can seriously affect the safety of the aircraft and the pilots.

On the other hand, in Korean Patent Laid-Open Publication No. 10-2010-0061946 ("Steering displacement sensor detection system for control input device for electronic flight control system", Patent document 1), two displacements of three displacement sensors mounted on the steering input device Discloses a system for detecting a normally operating displacement sensor in the event of a sensor failure. However, Patent Document 1 does not disclose any backup means that can replace the electronic flight control system in the event of a failure.

Korean Patent Publication No. 10-2010-0061946 (June 10, 2010)

SUMMARY OF THE INVENTION The present invention has been made to solve the above-mentioned problems, and it is an object of the present invention to provide an electronic flight control system capable of electronically controlling a flight body by driving a rotor operating part according to an operation part command generated by a flight control computer, And an object of the present invention is to provide a flight control system having backup means capable of controlling a flying object in the event of a failure in a control system.

According to an aspect of the present invention, there is provided an electronic flight control system having a mechanical steering apparatus as a backup, the electronic flight control system comprising: an electronic steering device for outputting an electric signal in response to an operation of a pilot; A sensor unit for outputting motion data according to the motion of the flying object; A rotor operating part for operating the rotor of the air vehicle; A flight control computer for generating an operating unit command using the electric signal output from the electronic control unit and the motion data output from the sensor unit; A driving unit for controlling the pitch axis, the roll axis, the collective axis, and the yaw axis of the flying object according to an operating unit command generated by the flight control computer to drive the rotor operating unit; And a mechanical steering device for controlling the pitch axis, the roll axis, the collective axis and the yaw axis of the air vehicle according to an operation of the pilot under the condition that the driving portion does not drive the rotor actuating portion to drive the rotor actuating portion .

Here, the driving unit may include two electric actuators provided on each of the pitch axis, the roll axis, the collective axis and the yaw axis of the air vehicle, and the electric actuator may be configured to change the displacement according to an operating unit command generated in the flight control computer The pitch axis, the roll axis, the collective axis and the yaw axis of the air vehicle are changed.

Here, if any one of the actuators of the two electric actuators provided on the pitch axis, the roll axis, the collective axis and the yaw axis of the air vehicle has a failure and is fixed at a specific position, By controlling the other electric actuator in the shaft provided with the electric actuator so that the sum of the displacements of the electric actuator in which the failure occurs and the electric actuator in the other is made to be zero so that the driving part does not drive the rotor actuating part And a condition is set.

Alternatively, when the flight control computer can not generate the operating unit command, the flight control computer generates and transmits a failure signal to the electric actuator, and when the electric actuator receives the failure signal, Is set to a neutral position in which the displacement of the rotor is fixed to zero so that the drive unit does not drive the rotor operating part.

Here, a failure occurs in any of the actuators of the two electric actuators provided on each of the pitch axis, the roll axis, the collective axis and the yaw axis of the air vehicle, so that the displacement of the electric actuator can not be fixed to the neutral position, The flight control computer controls the other electric actuator on the axis provided with the failed electric actuator so that the sum of the displacements of the failed electric actuator and the other electric actuator is set to 0 Thereby setting a condition that the driving unit does not drive the rotor operating unit.

The present invention relates to a control system for a hybrid vehicle, which transmits an operating unit command generated by a flight control computer to a driving unit connected to a pitch axis, a roll axis, a collective axis and a yaw axis of a flight body so that the driving unit drives the rotor operating unit, The pilot can mechanically drive the pitch axis, the roll axis, the collective axis and the yaw axis of the flight body directly through the mechanical steering device under the condition of not driving the part (in case of failure in the electronic flight control system) . That is, according to the present invention, even if a malfunction occurs in the electronic flight control system, the pilot can drive the rotor operating part by adjusting the pitch axis, the roll axis, the collective axis and the yaw axis of the air vehicle through the mechanical steering device Therefore, it is possible to secure the safety of aviation and pilots.

Further, according to the present invention, it is possible to easily convert a conventional mechanical flight control system into a flight control system according to the present invention without any modification of the front face. Therefore, it is possible to save time and cost required to implement the flight control system according to the present invention do.

1 is a schematic view of a conventional mechanical flight control system.
2 is a schematic view of a conventional electronic flight control system.
3 is a schematic diagram of a flight control system according to an embodiment of the present invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS An electronic flight control system having a mechanical steering apparatus according to the present invention as a backup will be described in detail with reference to the accompanying drawings. As used herein, unless otherwise defined, technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this invention belongs. In the following description and the accompanying drawings, The description of the known functions and configurations that may unnecessarily obscure the detailed description will be omitted. Also, in the accompanying drawings, a solid line between a component and a component means a connection through a mechanical linkage, and a dotted line between a component and a component means a connection through a wired or wireless communication.

Before describing the present invention in detail, a conventional mechanical flight control system and an electronic flight control system will be schematically described with reference to FIGS. 1 and 2. FIG.

FIG. 1 is a schematic view of a conventional mechanical flight control system. In the conventional mechanical flight control system, a pilot controls a cyclic stick 51, a collective stick 52, When the pedal 53 is operated, the control command is configured to drive the actuator unit 60 through the mechanical linkage.

Concretely, the cyclic stick 51 is connected to the pitch axis 1 and the roll axis 2 of the mechanical linkage, the collective stick 52 is connected to the collective shaft 3 of the mechanical linkage, (53) is connected to the yaw axis (4) of the mechanical linkage. Thus, when the pilot operates the cyclic stick 51, the collective stick 52 and the pedal 53, the pitch axis 1, the roll axis 2, the collective axis 3 and the yaw axis 4) are controlled so that the actuator unit 60 can be driven.

In the conventional mechanical flight control system, an electric actuator 40 is connected to the pitch axis 1, the roll axis 2, the collective axis 3 and the yaw axis 4, 50 to actuate the actuator unit 60 according to the steering command generated.

The electric actuator 40 receives the motion data of the airplane outputted from the sensor unit 20 and uses the motion data to control the operation of the pilot's mechanical manipulator 50 separately from the operation of the pilot's mechanical manipulator 50 The pitch axis 1, the roll axis 2, the collective axis 3 and the yaw axis 4 are adjusted. As the electric actuator 40 performs the control of each axis within the limited control range, the tremor generated in the mechanical control device 50 can be alleviated when the pilot controls the airplane, .

However, the control authority of the electric actuator 40 is very limited in the mechanical flight control system. That is, in the mechanical flight control system, since the pilot is designed so that the pilot manipulates the mechanical manipulator 50, the electric actuator 40 can be operated only by about 10% The actuator control unit 60 has a control authority to drive the actuator unit 60 separately.

FIG. 2 is a schematic view of a conventional electronic flight control system. In the conventional electronic flight control system, a pilot controls the cyclic steering stick assembly 11 'and the collective pilot When the stick assembly 12 'is operated, each assembly 11', 12 'generates and outputs an electric signal in response to the operation of the pilot.

Specifically, in the cyclic steering stick assembly 11 ', an electric signal relating to the pitch, roll and yaw of the air vehicle is outputted to the flight control computer 30' in response to the operation of the pilot, and the collective steering stick assembly 12 ' ) Outputs an electric signal relating to the collective of the flying object to the flight control computer 30 'in response to the manipulation of the pilot. At this time, the electronic control unit 10 'and the flight control computer 30' are connected by wire or wireless communication instead of a mechanical linkage.

The flight control computer 30 'receives electric signals relating to the pitch, roll, yaw and collective output from the electronic control unit 10' and also inputs motion data of the airplane outputted from the sensor unit 20 ' Receive. The flight control computer 30 'generates an operation unit command for driving the actuator unit 60' using the received electric signal and the motion data, and the actuator unit 60 ' So that the pilot can operate the rotor 70 'to control the pilot as desired.

In this way, in the electronic flight control system, the flight control computer 30 'has the control authority to completely drive the actuator unit 60'. Therefore, when a malfunction occurs in the electronic flight control system (specifically, when a malfunction occurs in the electronic control unit 10 ', the sensor unit 20' or the flight control computer 30 '), There is a problem in that the safety of the airplane and the pilots is difficult to be secured because there is no way to control the airplane.

Accordingly, in the present invention, when a failure occurs in the electronic flight control system, the pilot controls the pilot by using the electronic flight control system, We propose a flight control system.

3 is a schematic view of a flight control system according to an embodiment of the present invention. The flight control system according to an embodiment of the present invention includes an electronic control unit 100, a sensor unit 200, a flight control computer 300, a driving part 400, a mechanical steering device 500, and a rotor operating part 600.

The electronic steering apparatus 100 may include a cyclic steering stick assembly 110 and a collective steering stick assembly 120. When the pilot operates the electronic steering apparatus 100, And generates and outputs an electric signal.

Specifically, in the cyclic pilot stick assembly 110, an electric signal relating to the pitch, roll, and yaw of the flying object is output to the flight control computer 300 in response to the manipulation of the pilot. In the collective pilot stick assembly 120, And outputs an electric signal relating to the collective of the flying object to the flight control computer 300. At this time, the electric signal outputted from the electronic control unit 100 is transmitted to the flight control computer 300 through wired or wireless communication.

The sensor unit 200 senses various motions (e.g., angular velocity, attitude, altitude, etc.) of the flying object during flight, and outputs the motion data of the flying object. The sensor unit 200 may be composed of a plurality of sensors capable of detecting various motions of the air vehicle and converting the motion data into data. The motion data of the air vehicle outputted from the sensor unit 200 may be transmitted to the flight control computer 300 by wire or wireless Can be communicated via communication.

Meanwhile, in the present invention, the flying object means an object flying by the rotation of the rotor 700 like a helicopter, and the rotor 700 may be composed of a main rotor and a tail rotor.

The rotor 700 is operated by the rotor operating portion 600 and the rotor operating portion 600 includes three main rotor actuators (MRA) for controlling the operation of the main rotors and the operation of the tail rotor (Tail Rotor Actuator (TRA)).

The drive unit 400 and the rotor operation unit 600 may include a pitch axis 1000, a roll axis 2000, a collective axis 3000, And a yaw axis 4000, respectively. That is, the rotor operating part 600 is driven by the driving part 400 and the rotor 700 is operated (rotated) by the rotor operating part 600, .

The flight control computer 300 receives electric signals relating to the pitch, roll, yaw and collective output from the electronic control unit 100 and also receives motion data of the airplane output from the sensor unit 200.

Then, the flight control computer 300 generates an operation unit command for driving the rotor operation unit 600 using the received electric signal and the motion data, and the rotor operation unit 600 controls the operation of the rotor operation unit 600 according to the operation unit command By operating the rotor 700 of the airplane, the pilot can steer the airplane as desired.

Here, the operation unit command generated by the flight control computer 300 may be generated by applying the electric signal and the motion data to an automatic flight control algorithm stored in advance in the flight control computer 300. Since the operation unit command reflects not only the electrical signal generated by the pilot's operation of the electronic control unit 100 but also the steering compensation value corresponding to the motion data of the flying object, the disturbance due to the motion of the airplane can be canceled, Characteristics can be implemented, and pilot control of the pilot can be achieved.

The operating unit commands generated in the flight control computer 300 are transmitted to the driving unit 400 connected to the pitch axis 1000, the roll axis 2000, the collective axis 3000 and the yaw axis 4000 of the air vehicle. The driving unit 400 controls the pitch axis 1000, the roll axis 2000, the collective axis 3000 and the yaw axis 4000 of the air vehicle to drive the rotor operation unit 600 according to the operation unit command.

Since the driving unit 400 drives the rotor operation unit 60 according to the operation unit command generated in the flight control computer 300, unlike the electric actuator 40 used in the mechanical flight control system shown in FIG. 1, You must have control over the scope. That is, the driving unit 400 is designed to be operated in a range in which the pitch axis 1000, the roll axis 2000, the collective axis 3000, and the yaw axis 4000 of the air vehicle are adjusted in a fully displaceable range There is a need.

The driving unit 400 includes an electronic actuator (EA) having two units (eight units in total) provided on each of the pitch axis 1000, the roll axis 2000, the collective axis 3000 and the yaw axis 4000 of the air vehicle. Each of the electric actuators controls the pitch axis 1000, the roll axis 2000, the collective axis 3000, and the yaw axis 3000 of the flight body by adjusting the displacement according to an operation unit command generated in the flight control computer 300 4000).

At this time, the two electric actuators provided on any one axis of the pitch axis 1000, the roll axis 2000, the collective axis 3000 and the yaw axis 4000 of the air vehicle are connected to each other in series, It is preferable that the actuators are each configured to have a control authority of 50% each. For example, when the driving unit 400 rotates the pitch axis 1000 of the flying object upward (that is, in the direction of the driving unit 600 from the driving unit 400) in accordance with the operating unit command generated in the flight control computer 300 Each of the two electric actuators connected in series with each other on the pitch axis 1000 can be configured to be located at a position 10 mm away from the neutral position where the displacement is zero.

As described above, in the case where a failure occurs in any one of the two electric actuators provided on the respective axes (that is, when the actuator is fixed to a specific position Even when the displacement adjustment is impossible), it is possible to easily and quickly escape from the danger due to the failure.

In addition, by transmitting the operating unit command generated in the flight control computer 300 to the driving unit 400 and driving the driving unit 400 to drive the rotor operating unit 600, Therefore, it is possible to reduce the cost and time required to implement the flight control system according to the present invention, and to provide a completely new electronic flight control system It is possible to minimize the risk that may arise in the case of development.

As described above, when a malfunction occurs in the electronic flight control system, it is required that the pilot and the pilot are secured at the same time by providing a backup device as an alternative to the pilot. To this end, the present invention proposes a method of providing a mechanical control device 500 of a mechanical flight control system as a backup control device of an electronic flight control system.

The mechanical steering device 500 comprises a cyclic stick 510, a collective stick 520 and a pedal 530, wherein the cyclic stick 510 comprises a pitch axis 1000 and a roll axis 2000 of a mechanical linkage, The collective stick 520 is connected to the collective axis 3000 of the mechanical linkage and the pedal 530 is connected to the yaw axis 4000 of the mechanical linkage.

However, in the conventional mechanical flight control system, as described above, the electric actuator 40 is configured to have only a limited limited authority. However, in the present invention, the driving unit 400 includes the operation Since the subcommand must be executed as it is, it is configured to have full scope of authority (Fully Authority).

However, in the case where a malfunction occurs in the electric actuator such as the case where the electric actuator constituting the driving unit 400 is fixed at a specific position and the displacement can not be adjusted, or the electric actuator 100, the sensor unit 200, When a failure occurs in the flight control computer 300 and the flight control computer can not generate an operation unit command, the drive amount of the drive unit 400 is made zero (more specifically, The sum of the displacements of the actuators is made to be zero), the driver 400 sets the conditions for not driving the rotor actuating part 600, so that the pilot sets the mechanical actuating device 500 It is desirable to be able to steer the air vehicle through.

Hereinafter, an embodiment in which the driving unit 400 does not drive the rotor operation unit 600 will be described. The electronic control unit 100, the sensor unit 200, the flight control computer 300 Or when a failure occurs in the driving unit 400, the pilot can cope safely and quickly through the mechanical control unit 500. [

As described above, two electric actuators may be provided for each of the pitch axis 1000, the roll axis 2000, the collective axis 3000 and the yaw axis 4000 of the air vehicle. One of the electric actuators For example, when a failure occurs in any one of the two electric actuators in the roll axis 2000 and the displacement is not adjustable while being fixed at a specific position, It is not possible to perform the axis adjustment (for example, the roll axis 2000 adjustment) according to the operation unit command generated in the computer 300, so that the entire driving unit 400 should not drive the rotor operation unit 600 .

The flight control computer 300 and the electric actuator can be connected by wire or wireless communication so that the flight control computer 300 can detect the occurrence of a failure of the electric actuator and control the operation of the electric actuator.

For example, if the flight control computer 300 determines that one of the two electrical actuators in the roll axis 2000 is in a particular position (e.g., a position at -10 mm relative to a neutral position with a displacement of zero) If it is detected that there is no movement while being fixed, the other electric actuator in the roll axis 2000 is controlled to make the sum of displacements of the electric actuator in which the failure occurs and the other electric actuator to zero For example, the displacement of the other electric actuator is controlled to a position of 10 mm with respect to the neutral position), thereby setting a condition that the driving unit 400 does not drive the rotor operation unit 600. At the same time, the flight control computer 300 controls the two electric actuators in the pitch axis 1000, the collective axis 3000 and the yaw axis 4000 so that all displacements of the electric actuators are fixed at the neutral position (That is, making the sum of the displacements of the two electric actuators in each axis all zero), the condition that the driving section 400 does not drive the rotor operating section 600 can be set.

When the condition that the driving unit 400 does not drive the rotor operation unit 600 is set as described above, the flight control computer 300 stops generating the operation unit command further, Only mechanical control becomes possible. That is, the pilot can control the pitch axis 1000, the roll axis 2000, the collective axis 3000 and the yaw axis 4000 of the air vehicle by operating the mechanical steering device 500 instead of the electronic steering device 100, In this case, the pitch axis 1000, the roll axis 2000, the collective axis 3000, and the yaw axis 4000, rather than the driving unit 400, directly drive the rotor operation unit 600.

When the communication between the flight control computer 300 and the electronic control unit 100 is disconnected and the communication between the flight control computer 300 and the sensor unit 200 is disconnected, the electronic control unit 100, the sensor unit 200 When the flight control computer 300 can not generate an operating unit command, such as when a problem occurs in the flight control computer 300 itself or when a problem occurs in the flight control computer 300 itself, The failure signal can be generated and the failure signal can be transmitted to each of the electric actuators constituting the driving unit 400.

When the electric actuator receives the failure signal, the electric actuator may be set to a condition that the actuator 400 is not driven by the rotor 400 by being fixed at the neutral position where the displacement is zero.

At this time, if the flight control computer 300 determines that any one of the two electric actuators in the roll axis 2000 is located at a specific position (for example, at a position of -10 mm relative to the neutral position with a displacement of 0 , It controls the other electric actuator in the roll axis 2000 so that the sum of the displacements of the electric actuator in which the failure occurs and the other electric actuator is set to 0 (For example, by controlling the displacement of the other electric actuator to a position of 10 mm with respect to the neutral position), it is possible to set a condition that the driving unit 400 does not drive the rotor operation unit 600.

When the condition that the driving unit 400 does not drive the rotor operation unit 600 is set as described above, the flight control computer 300 stops generating the operation unit command further, Only mechanical control becomes possible. That is, the pilot can control the pitch axis 1000, the roll axis 2000, the collective axis 3000 and the yaw axis 4000 of the air vehicle by operating the mechanical steering device 500 instead of the electronic steering device 100, In this case, the pitch axis 1000, the roll axis 2000, the collective axis 3000, and the yaw axis 4000, rather than the driving unit 400, directly drive the rotor operation unit 600.

As described above, according to the present invention, even if a malfunction occurs in the electronic flight control system, the pilot can control the pitch axis 1000, the roll axis 2000, the collective axis 3000, Since the yaw axis 4000 is controlled to drive the rotor operation unit 600, it is possible to secure the safety of the flying body and the pilot.

It should be understood that the above-described embodiments are illustrative in all aspects and not restrictive. The scope of the present invention is defined by the appended claims rather than the description of the invention, and all changes or modifications derived from the meaning and scope of the claims and their equivalents are to be construed as being included within the scope of the present invention do.

100: Electronic control device
200:
300: flight control computer
400:
500: Mechanical controls
600: rotor operating part
700: Rotor
1000: pitch axis
2000: Roll axis
3000: Collective axis
4000: Yodo

Claims (5)

An electronic steering device for outputting an electric signal in response to an operation of a pilot;
A sensor unit for outputting motion data according to the motion of the flying object;
A rotor operating part for operating the rotor of the air vehicle;
A flight control computer for generating an operating unit command using the electric signal output from the electronic control unit and the motion data output from the sensor unit;
A driving unit for controlling the pitch axis, the roll axis, the collective axis, and the yaw axis of the flying object according to an operating unit command generated by the flight control computer to drive the rotor operating unit; And
And a mechanical control device for controlling the pitch axis, the roll axis, the collective axis and the yaw axis of the air vehicle according to the manipulation of the pilot under the condition that the driving portion does not drive the rotor actuating portion to drive the rotor actuating portion,
Wherein the driving unit comprises two electric actuators each including two pitch axes, a roll axis, a collective axis and a yaw axis of the air vehicle,
The electric actuator controls a pitch axis, a roll axis, a collective axis and a yaw axis of the airplane by changing a displacement according to an operation unit command generated in the flight control computer,
When a failure occurs in any one of the electric actuators provided on the pitch axis, the roll axis, the collective axis and the yaw axis of the air vehicle,
Wherein the flight control computer controls another electric actuator on the axis provided with the failed electric actuator to make the sum of the displacements of the failed electric actuator and the other electric actuator zero, Wherein the electronic control unit sets a condition that the driving unit does not drive the rotor operating unit.
delete delete An electronic steering device for outputting an electric signal in response to an operation of a pilot;
A sensor unit for outputting motion data according to the motion of the flying object;
A rotor operating part for operating the rotor of the air vehicle;
A flight control computer for generating an operating unit command using the electric signal output from the electronic control unit and the motion data output from the sensor unit;
A driving unit for controlling the pitch axis, the roll axis, the collective axis, and the yaw axis of the flying object according to an operating unit command generated by the flight control computer to drive the rotor operating unit; And
And a mechanical steering device for controlling the pitch axis, the roll axis, the collective axis and the yaw axis of the air vehicle in accordance with the operation of the pilot under the condition that the driving portion does not drive the rotor operating portion,
Wherein the driving unit comprises two electric actuators each including two pitch axes, a roll axis, a collective axis and a yaw axis of the air vehicle,
The electric actuator controls a pitch axis, a roll axis, a collective axis and a yaw axis of the airplane by changing a displacement according to an operation unit command generated in the flight control computer,
When the flight control computer can not generate the operating unit command, the flight control computer generates and transmits a failure signal to the electric actuator,
Characterized in that when the electric actuator receives the fault signal, a condition is established in which the drive unit is not driven by the electric actuator being fixed at a neutral position where the displacement of the electric actuator is zero. The electronic flight control system.
5. The method of claim 4,
A failure occurs in any one of the electric actuators provided on each of the pitch axis, the roll axis, the collective axis and the yaw axis of the air vehicle, so that the displacement of the electric actuator can not be fixed to the neutral position, If so,
Wherein the flight control computer controls another electric actuator on the axis provided with the failed electric actuator to make the sum of the displacements of the failed electric actuator and the other electric actuator zero, Wherein the electronic control unit sets a condition that the driving unit does not drive the rotor operating unit.

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