Hybrid control configuration method applied to high-maneuverability micro-nano satellite

Abstract

The invention discloses a hybrid control configuration method applied to a high-maneuverability micro-nano satellite, which takes three reaction flywheels and a pair of control moment gyroscopes as a hybrid control mechanism, so that the control moment gyroscopes give consideration to the backup function of the three reaction flywheels while exerting high maneuverability in a single axis, and the reliability of a satellite attitude control system is improved.

Description

Technical Field

[0001] The present invention belongs to the technical field of satellite attitude and orbit control, and particularly relates to a hybrid control mechanism configuration method applicable to a single-axis highly maneuverable satellite. Background Art

[0002] With the continuous development of satellite research and development technologies, the requirements for the high-speed maneuverability of satellites are also increasing. Traditional satellites usually use reaction wheels, magnetorquers, and jet assemblies for attitude control. Among them, the magnetorquer adjusts the angular velocity of the satellite body according to the different magnetic field environments where the satellite is located; the jet assembly sprays substances through nozzles to control the attitude of the satellite; the reaction wheel exchanges angular momentum with the satellite body through a high-speed rotor to achieve the attitude control of the satellite, and the reaction wheel is the main component for satellite attitude control.

[0003] Currently, due to the limitations of reaction wheel bearings and materials, if the angular momentum of the reaction wheel is to be increased, its weight and size will also increase significantly. For micro-nano satellites with high requirements for maneuverability, the use of reaction wheels cannot guarantee the overall satellite weight under high maneuverability. Although the jet assembly can improve the attitude control efficiency, it consumes fuel, and in the case of limited fuel reserves, long-term attitude control of the satellite cannot be achieved. As a spacecraft attitude control component, the control moment gyro has excellent torque output capabilities because its output torque acts directly on the satellite body through bearings, and the torque transmission is independent of the torque device of the frame servo system. It does not involve fuel consumption, and its maneuverability is significantly better than that of reaction wheels under the same weight. However, it is expensive and occupies a relatively large space in layout, and is usually not suitable for small micro-nano satellites.

[0004] Patent CN201410163314.3 proposes a satellite steady-state control method using a control moment gyro to replace the momentum wheel. It is proposed to use the high-speed rotor of the backup control moment gyro as a fixed angular momentum momentum wheel, and provide a fixed bias angular momentum in a certain direction by rotating the low-speed frame angle, so that the momentum wheel can work in a biased state, avoiding the accuracy loss problem of using the momentum wheel when it passes through zero. This method enables the cold backup control moment gyro to serve as a backup for both side-sway maneuvers and steady-state control, improving the reliability of the actuator in orbit. However, the control moment gyro is only used as a backup flywheel and cannot give full play to its advantages of large torque and fast response. Patent CN201910960844.3 proposes a method for realizing three-axis stable control of a satellite with a single control moment gyro and two flywheels, which can achieve attitude stable control when only one control moment gyro and two reaction wheels are effective on the satellite, but cannot improve the maneuverability of the satellite. Summary of the Invention

[0005] The object of the present invention is to overcome the limitation of the maneuverability of existing microsatellites, and provide a hybrid control configuration method for high-maneuverability microsatellites. Three reaction wheels and a pair of control moment gyroscopes are used as a hybrid control mechanism, so that the control moment gyroscopes can exert high maneuverability in a single axis while taking into account the backup function of the three reaction wheels, and improving the reliability of the satellite attitude control system.

[0006] The technical solution of the present invention is: a hybrid control configuration method for high-maneuverability microsatellites. Three reaction wheels and a pair of control moment gyroscopes are used as a hybrid control mechanism, so that the control moment gyroscopes can exert high maneuverability in a single axis while taking into account the backup function of the three reaction wheels, and improving the reliability of the satellite attitude control system.

[0007] Furthermore, the specific steps are as follows:

[0008] Step 1: Arrange three reaction wheels on the satellite; the reaction wheel group adopts a three-inclined installation configuration, and the angular momentum direction of a single reaction wheel is inclined to the three axes of the satellite body by ;

[0009] Step 2: Arrange a group of control moment gyroscopes on the satellite; determine the satellite body axis that needs to perform high maneuvering according to the mission requirements, and arrange the frame axis and the initial angular momentum direction of the control moment gyroscopes; the frame axis direction is not the direction of the torque output of the control moment gyroscopes, and the angular momentum directions of the two control moment gyroscope groups are kept opposite, which can avoid the angular momentum interference caused by the on-orbit spin-up of the control moment gyroscopes;

[0010] Step 3: According to the hybrid control configuration of the reaction wheel group and the control moment gyroscopes, realize the three-axis control of the satellite; use the control strategy of three-wheel control to perform the three-axis steady-state control of the satellite in the steady-state mission; in the high-maneuver mission, use the control moment gyroscope group to realize the rapid maneuvering process of a single axis;

[0011] Step 4: When any reaction wheel fails and the satellite cannot achieve three-axis steady-state control, switch the control logic, lock one of the control moment gyroscope groups in the low-speed frame axis state, and use the variable speed of the high-speed rotor part as a backup reaction wheel to maintain the three-axis steady-state control of the satellite; when the control moment gyroscope is used as a backup reaction wheel, the satellite will only be controlled by the reaction wheel group.

[0012] Furthermore, in Step 1, the installation matrix of the reaction wheel group is:

[0013] 。

[0014] Furthermore, in Step 2, the installation matrix of the control moment gyroscope group is:

[0015]

[0016] Among them, matrix A is the direction of the angular momentum vector after the control moment gyroscope frame angle rotates by 90°, and matrix B is the direction of the angular momentum vector when the control moment gyroscope frame angle is 0°.

[0017] Furthermore, in step four, to ensure that the components of the high-speed component of the control moment gyroscope in the satellite body axis direction are equal, the direction of the angular momentum of the high-speed rotor of the control moment gyroscope and the oblique installation angle with the corresponding satellite body axis are both:

[0018] .

[0019] The beneficial effects of the present invention are as follows: A hybrid control mechanism configuration method for high-maneuver micro-nano satellites is provided. By using three reaction wheels and a pair of control moment gyroscopes, the satellite can achieve three-axis high-precision steady-state control using three-wheel control under steady-state tasks, and can achieve the high-maneuverability of the satellite's single axis using a pair of control moment gyroscopes under high-maneuver tasks. At the same time, a pair of control moment gyroscopes can also be used as a backup means for reaction wheels to improve the reliability of the satellite system. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 It is a hybrid control configuration of a control moment gyroscope and a reaction wheel;

[0021] Figure 2 It is the first configuration of the control moment gyroscope switching to backup the reaction wheel;

[0022] Figure 3 It is the second configuration of the control moment gyroscope switching to backup the reaction wheel. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0023] The following further describes the present invention with reference to the accompanying drawings.

[0024] This embodiment provides a hybrid control configuration method for high-maneuver micro-nano satellites, using three reaction flywheels and a pair of control moment gyroscopes as a hybrid control mechanism, enabling the control moment gyroscope to exert high-maneuverability on a single axis while taking into account the backup function of the three reaction wheels, and improving the reliability of the satellite attitude control system.

[0025] Since the high-speed rotor part of the control moment gyroscope can only rotate along the frame axis, it can only provide control torques in two axial directions. Therefore, the reaction wheel group adopts a three-oblique installation configuration, and each reaction wheel forms a certain angle with the three axes of the satellite body. To ensure that the components of the angular momentum of the reaction wheel in the three axes of the satellite body are equal, that is:

[0026]

[0027] Among them, h1, h2, and h3 are the components of the reaction wheel angular momentum on the three axes of the spacecraft, and h is the reaction wheel angular momentum. Therefore, the direction of the reaction wheel angular momentum forms an angle with the skew angles of the three axes of the spacecraft as follows:

[0028]

[0029] At the same time, to ensure the high maneuverability of the satellite about a certain axis of the spacecraft, taking the high maneuverability about the pitch Y-axis of the spacecraft as an example, the control moment gyroscope gimbal axis and the angular momentum direction are distributed on the X-axis and Z-axis of the spacecraft. Taking the gimbal axis as the -X-axis of the spacecraft as an example, the angular momentum directions of the control moment gyroscopes are along the ±Z-axes of the spacecraft respectively. The arrangement with opposite angular momentum directions can cancel out the angular momenta of the high-speed rotors of the two control moment gyroscopes when the high-speed rotors of the control moment gyroscopes are pulled up, without generating additional disturbing angular momentum.

[0030] From this, the installation matrices of the reaction wheel group and the control moment gyroscope group in this configuration are as Figure 1 shown, realizing the closed-loop control of the three-axis attitude of the satellite. To prevent the reaction wheel group from having a singular solution, it is necessary to ensure that the matrix is not full rank. The installation matrix of the reaction wheel group is shown as follows:

[0031]

[0032] The installation matrix of the control moment gyroscope group is shown as follows:

[0033]

[0034] Among them, the A matrix is the direction of the angular momentum vector after the control moment gyroscope gimbal angle rotates by 90°, and the B matrix is the direction of the angular momentum vector when the control moment gyroscope gimbal angle is 0°.

[0035] When a certain reaction wheel of the satellite fails, the satellite cannot achieve three-axis steady-state control. At this time, one of the control moment gyroscope groups is kept in the locked state of the low-speed gimbal axis, and the high-speed rotor part is used as a backup reaction wheel by varying the speed to maintain the three-axis steady-state control of the satellite. To ensure that the components of the high-speed part of the control moment gyroscope in the axial direction of the spacecraft are equal, the direction of the angular momentum of the high-speed rotor of the control moment gyroscope forms an angle with the skew angle of the corresponding spacecraft axis as follows:

[0036]

[0037] When the control moment gyroscope is used as a backup reaction wheel, its hybrid control configuration is as Figure 2 , Figure 3 shown, and at this time the satellite will be controlled only by the reaction wheel group.

[0038] Step 1: Arrange three reaction wheels on the satellite. The reaction wheel assembly adopts a three-inclined configuration, and the angular momentum direction of a single reaction wheel forms an angle of with the three-axis inclined angles of the satellite body. To avoid singular values in the three-body control of the satellite and keep the installation matrix of the reaction wheel assembly full rank, determine the installation directions of the three reaction wheels. The installation method of the reaction wheel assembly can be appropriately modified according to the satellite structure. Taking one configuration as an example, the installation method of the reaction wheels is shown as follows:

[0039]

[0040] Step 2: Arrange a set of control moment gyros on the satellite. Determine the satellite body axis that needs to perform high maneuvers according to the mission requirements, and arrange the gimbal axis and the initial angular momentum direction of the control moment gyros. The gimbal axis direction is not the direction of the torque output of the control moment gyro, and the angular momentum directions of the two control moment gyro assemblies are kept opposite to avoid angular momentum interference caused by the on-orbit spin-up control moment gyro. Taking a micro-nano satellite with high maneuverability around the pitch axis as an example, the installation arrangement of the control moment gyros is shown as follows:

[0041]

[0042] Among them, matrix A is the angular momentum vector direction after the gimbal angle of the control moment gyro rotates 90°, and matrix B is the angular momentum vector direction when the gimbal angle of the control moment gyro is 0°.

[0043] Step 3: According to the hybrid control configuration of the reaction wheel assembly and the control moment gyros, achieve the three-axis control of the satellite. Use the control strategy of three-wheel control for the three-axis steady-state control of the satellite in the steady-state mission. In the high-maneuver mission, use the control moment gyro assembly to achieve the rapid maneuvering process of a single axis.

[0044] Step 4: When any reaction wheel fails and the satellite cannot achieve three-axis steady-state control, switch the control logic to lock the low-speed gimbal axis of the control moment gyro, and use the high-speed rotor as a backup reaction wheel to continue the steady-state control of the satellite. Taking the above configuration as an example, when the reaction wheel is unavailable, its command distribution matrix is adjusted accordingly to continue to achieve the three-axis control of the satellite:

[0045] If reaction wheel A is unavailable, then:

[0046]

[0047] If reaction wheel B is unavailable, then:

[0048]

[0049] If reaction wheel C is unavailable, then:

[0050]

[0051] Among them, the corresponding relationships of the matrix rows are Reaction Wheel A, Reaction Wheel B, Reaction Wheel C, Control Moment Gyro A, and Control Moment Gyro B respectively.

[0052] The above are only the preferred embodiments of the present invention. It should be pointed out that for those of ordinary skill in the art, without departing from the principle of the present invention, several improvements and modifications can be made, and these improvements and modifications should also be regarded as the protection scope of the present invention.

Claims

1. A hybrid control configuration method applied to a highly maneuverable micro-nano satellite, characterized in that: Three reaction flywheels and a pair of control moment gyro are used as a hybrid control mechanism, so that the control moment gyro can exert high maneuverability on a single axis while taking into account the backup function of the three reaction wheels, thereby improving the reliability of the satellite attitude control system.

2. The hybrid control configuration method for a highly maneuverable micro-nano satellite according to claim 1, characterized in that: The specific steps are as follows: Step 1: Arrange three reaction wheels on the satellite; the reaction wheel group adopts a three-slant configuration, and the angular momentum direction of a single reaction wheel is in the same direction as the three-axis slant angle of the satellite. ; Step 2: Arrange a group of control moment gyros on the satellite; determine the satellite axis that needs to be highly maneuverable according to the mission requirements, and arrange the frame axis and initial direction of angular momentum of the control moment gyro; the frame axis direction is not the direction of the control moment gyro torque output, and the angular momentum directions of the two control moment gyro groups remain in opposite states, which can avoid angular momentum interference caused by the control moment gyro spinning on orbit; Step 3: According to the hybrid control configuration of the reaction wheel group and the control moment gyro, the three-axis control of the satellite is realized; in the steady-state mission, the control strategy of the three-wheel control is used to perform the steady-state control of the three axes of the satellite; in the high-maneuverability mission, the control moment gyro group is used to realize the rapid maneuvering process of a single axis; Step 4: When any reaction wheel fails and the satellite cannot achieve three-axis steady-state control, the control logic is switched to keep one of the control torque gyro groups in a low-speed frame shaft locked state, and the high-speed rotor part is used as a backup reaction wheel by changing the speed to maintain the satellite's three-axis steady-state control; when the control torque gyro is used as a backup reaction wheel, the satellite will only be controlled by the reaction wheel group.

3. The hybrid control configuration method for a highly maneuverable micro-nano satellite according to claim 2, characterized in that: In step 1, the installation matrix of the reaction wheel set is: 。 4. The hybrid control configuration method for a highly maneuverable micro-nano satellite according to claim 2, characterized in that: In step 2, the installation matrix of the control moment gyro group is: Among them, the A matrix is ​​the direction of the angular momentum vector after the control moment gyro frame angle rotates 90°, and the B matrix is ​​the direction of the angular momentum vector when the control moment gyro frame angle is 0°.

5. The hybrid control configuration method for a highly maneuverable micro-nano satellite according to claim 2, characterized in that: In step 4, in order to ensure that the components of the high-speed components of the control moment gyro are equal on the axis of the star body, the angular momentum direction of the high-speed rotor of the control moment gyro and the oblique angle of the corresponding star body axis are both: 。

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