A Hybrid Control Configuration Method for Highly Maneuverable Micro / Nano Satellites

Abstract

This invention discloses a hybrid control configuration method for highly maneuverable micro-nano satellites, which uses three reaction flywheels and a pair of control moment gyroscopes as a hybrid control mechanism. This allows the control moment gyroscopes to exert high maneuverability on a single axis while also serving as backup for the three reaction wheels, thereby improving the reliability of the satellite attitude control system.

Description

Technical Field

[0001] This invention belongs to the field of satellite attitude and orbit control technology, and in particular, it is a hybrid control mechanism configuration method suitable for single-axis high-maneuverability satellites. Background Technology

[0002] With the continuous development of satellite development technology, the requirements for high-speed maneuverability of satellites are also constantly increasing. Traditional satellites typically use reaction wheels, magnetic torquers, and jet propulsion systems for attitude control. Among them, the magnetic torquer adjusts the angular velocity of the satellite according to different magnetic field environments; the jet propulsion system controls the satellite's attitude by ejecting material through nozzles; and the reaction wheel achieves attitude control by exchanging angular momentum with the satellite through a high-speed rotor. The reaction wheel is the main component for satellite attitude control.

[0003] Currently, due to limitations in bearings and materials for reaction wheel flywheels, increasing their angular momentum significantly increases their weight and size. For micro / nano satellites requiring high maneuverability, the use of reaction wheels cannot guarantee the overall satellite weight under high maneuverability. While jet propulsion systems can improve attitude control efficiency, they consume fuel, making long-term attitude control impossible with limited fuel reserves. Control moment gyroscopes, as a spacecraft attitude control component, directly transmit torque to the satellite through bearings, independent of the torque generator in the frame servo system. They possess excellent torque output capability, do not involve fuel consumption, and offer significantly better maneuverability than reaction wheels for the same weight. However, they are expensive and space-consuming, generally unsuitable for small micro / nano satellites.

[0004] Patent application CN201410163314.3 proposes a satellite steady-state control method using a control moment gyroscope instead of a momentum wheel. It proposes using a backup control moment gyroscope high-speed rotor as a fixed angular momentum wheel, providing a fixed bias angular momentum in a specific direction by rotating a low-speed frame angle, allowing the momentum wheel to operate in an biased state and avoiding the accuracy loss problem caused by the momentum wheel crossing zero. This method allows the cold backup control moment gyroscope to serve as a backup for both yaw maneuvers and steady-state control, improving the reliability of the actuator in orbit. However, since the control moment gyroscope is only used as a backup flywheel, its advantages of high torque and fast response cannot be fully utilized. Patent application CN201910960844.3 proposes a method for achieving three-axis stable control of a satellite using a single control moment gyroscope and two flywheels. This method can achieve attitude stability control when only one control moment gyroscope and two reaction flywheels are effective on the satellite, but it cannot improve the satellite's maneuverability. Summary of the Invention

[0005] The purpose of this invention is to overcome the limited maneuverability of existing micro and nano satellites and provide a hybrid control configuration method for highly maneuverable micro and nano satellites. This method uses three reaction flywheels and a pair of control moment gyroscopes as a hybrid control mechanism, so that the control moment gyroscopes can exert high maneuverability on a single axis while also taking into account the backup function of the three reaction wheels, thereby improving the reliability of the satellite attitude control system.

[0006] The technical solution of this invention is: a hybrid control configuration method for highly maneuverable micro-nano satellites, which uses three reaction flywheels and a pair of control moment gyroscopes as a hybrid control mechanism, so that the control moment gyroscopes can exert high maneuverability on a single axis while also taking into account the backup function of the three reaction wheels, thereby 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 assembly adopts a three-oblique mounting configuration, with the angular momentum direction of each individual reaction wheel perpendicular to the oblique mounting angle of the satellite's three axes.

[0009] Step 2: Arrange a set of control moment gyroscopes for the satellite; determine the satellite axis that requires high maneuverability according to the mission requirements, and arrange the frame axis and 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 gyroscope, and the angular momentum directions of the two control moment gyroscope groups are kept in opposite states to avoid angular momentum interference caused by the control moment gyroscopes spinning in orbit;

[0010] Step 3: Based on the hybrid control configuration of reaction wheel assembly and control moment gyroscope, realize the three-axis control of the satellite; in steady-state missions, use the three-wheel control strategy to achieve steady-state control of the satellite's three axes; in high-maneuver missions, use the control moment gyroscope assembly to achieve rapid maneuvering of a single axis.

[0011] Step 4: When any reaction wheel malfunctions and causes the satellite to be unable to achieve three-axis steady-state control, switch the control logic, keep one of the control moment gyroscopes in a low-speed frame axis locked, and use the high-speed rotor part as a backup reaction wheel by changing its speed to maintain the satellite's three-axis steady-state control; 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 one, the installation matrix of the reaction wheel assembly is as follows:

[0013]

[0014] Furthermore, in step two, the installation matrix for the control torque gyroscope assembly is as follows:

[0015]

[0016] In this matrix, matrix A represents the direction of the angular momentum vector after the control moment gyroscope frame angle has rotated 90°, and matrix B represents 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 control torque gyroscope's components along the celestial axis are equal, the direction of the high-speed rotor angular momentum of the control torque gyroscope is made to be at the same angle as the corresponding oblique angle of the celestial axis:

[0018]

[0019] The beneficial effects of this invention are: it provides a hybrid control mechanism configuration method for highly maneuverable micro / nano satellites. By using three reaction wheels and a pair of control moment gyroscopes, the satellite can achieve high-precision three-axis steady-state control using the three-wheel control system under steady-state missions, and achieve high single-axis maneuverability using the pair of control moment gyroscopes under highly maneuverable missions. Simultaneously, the pair of control moment gyroscopes can also serve as a backup for the reaction wheels, improving the reliability of the satellite system. Attached Figure Description

[0020] Figure 1 A hybrid control configuration combining a torque gyroscope and a reaction wheel;

[0021] Figure 2 To control the torque gyroscope, switch to backup reaction wheel configuration one;

[0022] Figure 3 To control the torque gyroscope, switch to backup reaction wheel configuration two. Detailed Implementation

[0023] The present invention will now be further described with reference to the accompanying drawings.

[0024] This embodiment provides a hybrid control configuration method for highly maneuverable micro / nano satellites, which uses three reaction flywheels and a pair of control moment gyroscopes as a hybrid control mechanism. This allows the control moment gyroscopes to exert high maneuverability on a single axis while also providing backup functionality for the three reaction flywheels, thereby improving the reliability of the satellite attitude control system.

[0025] Since the high-speed rotor of the control torque gyroscope can only rotate along the frame axis, it can only provide control torque in two axes. Therefore, the reaction wheel assembly adopts a three-oblique configuration, with each reaction wheel forming a certain angle with the three axes of the gyroscope. To ensure that the angular momentum components of the reaction wheels are equal along the three axes of the gyroscope, i.e.:

[0026]

[0027] Where h1, h2, and h3 are the components of the reaction wheel's angular momentum along the three axes of the celestial body, and h is the angular momentum of the reaction wheel. Therefore, the direction of the reaction wheel's angular momentum is such that its angle with the oblique angle of the three axes of the celestial body is:

[0028]

[0029] Meanwhile, to ensure the satellite's high maneuverability on a certain axis of the celestial body, taking the high maneuverability of the celestial body's pitch Y-axis as an example, the control moment gyroscope frame axis and angular momentum direction are distributed on the celestial body's X-axis and Z-axis. Taking the frame axis as the celestial body's X-axis as an example, the control moment gyroscope angular momentum directions are respectively along the celestial body's ±Z-axis. The arrangement of opposite angular momentum directions can ensure that when the high-speed rotor of the control moment gyroscope is pulled up, the high-speed rotor angular momentum of the two control moment gyroscopes cancels each other out, without generating additional interfering angular momentum.

[0030] Therefore, the mounting matrix of the reaction wheel assembly and the control torque gyroscope group under this configuration can be obtained as follows: Figure 1 As shown, closed-loop control of the satellite's three-axis attitude is achieved. To prevent singular solutions from being generated by the reaction wheel assembly, the matrix must not be full rank. The installation matrix of the reaction wheel assembly is shown in the following equation:

[0031]

[0032] The mounting matrix for the control torque gyroscope assembly is shown in the following formula:

[0033]

[0034] In this matrix, matrix A represents the direction of the angular momentum vector after the control moment gyroscope frame angle has rotated 90°, and matrix B represents the direction of the angular momentum vector when the control moment gyroscope frame angle is 0°.

[0035] When a reaction wheel on the satellite malfunctions, the satellite cannot achieve three-axis steady-state control. In this case, one of the control moment gyroscopes is locked at a low speed on the frame axis, while the high-speed rotor acts as a backup reaction wheel by varying its speed, thus maintaining the satellite's three-axis steady-state control. To ensure that the components of the high-speed component of the control moment gyroscope are equal along the satellite's axis, the angular momentum direction of the high-speed rotor of the control moment gyroscope is angled with respect to the corresponding oblique angle of the satellite axis.

[0036]

[0037] When the control torque gyroscope is used as a backup reaction wheel, its hybrid control configuration is as follows: Figure 2 , Figure 3 As shown, at this point, the satellite will only be controlled by the reaction wheel assembly.

[0038] Step 1: Arrange three reaction wheels on the satellite. The reaction wheel assembly adopts a three-oblique mounting configuration, with the angular momentum direction of each individual reaction wheel perpendicular to the oblique angle of the satellite's three axes. To avoid singular values ​​in the satellite's three-body control and maintain the full rank of the reaction wheel assembly matrix, the installation orientation of the three reaction wheels is determined. The reaction wheel assembly installation method can be appropriately modified according to the satellite structure. The following example illustrates one configuration, with the reaction wheel installation method shown in the following equation:

[0039]

[0040] Step Two: Arrange a set of control moment gyroscopes for the satellite. Determine the satellite's required high-maneuverability axis based on mission requirements, and arrange the frame axis and initial angular momentum direction of the control moment gyroscopes accordingly. The frame axis direction is not the direction of the control moment gyroscope's torque output, and the angular momentum directions of the two control moment gyroscope sets are kept opposite to avoid angular momentum interference caused by the on-orbit spinning control moment gyroscopes. Taking a micro / nano satellite achieving high pitch axis maneuverability as an example, the installation arrangement of the control moment gyroscopes is shown in the following equation:

[0041]

[0042] In this matrix, matrix A represents the direction of the angular momentum vector after the control moment gyroscope frame angle has rotated 90°, and matrix B represents the direction of the angular momentum vector when the control moment gyroscope frame angle is 0°.

[0043] Step 3: Based on the hybrid control configuration of reaction wheel assembly and control moment gyroscope, realize the three-axis control of the satellite. In steady-state missions, the three-wheel control strategy is used to achieve steady-state control of the satellite's three axes. In high-maneuver missions, the control moment gyroscope assembly is used to achieve rapid maneuvering of a single axis.

[0044] Step 4: When any reaction wheel malfunctions, preventing the satellite from achieving three-axis steady-state control, the control logic is switched, locking the low-speed frame axis of the control moment gyroscope, while the high-speed rotor acts as a backup reaction wheel to continue the satellite's steady-state control. Taking the above configuration as an example, when the reaction wheel is unavailable, its command allocation matrix is ​​adjusted accordingly to continue achieving three-axis satellite control.

[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] The matrix rows correspond to reaction wheel A, reaction wheel B, reaction wheel C, control moment gyroscope A, and control moment gyroscope B, respectively.

[0052] The above description is only a preferred embodiment of the present invention. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the principle of the present invention, and these improvements and modifications should also be considered within the scope of protection of the present invention.

Claims

1. A hybrid control configuration method for highly maneuverable micro / nano satellites, characterized in that: Using three reaction flywheels and a pair of control moment gyroscopes as a hybrid control mechanism, the control moment gyroscopes can exert high maneuverability on a single axis while also serving as backup for the three reaction wheels, thereby improving the reliability of the satellite attitude control system. The specific steps are as follows: Step 1: Arrange three reaction wheels on the satellite; the reaction wheel assembly adopts a three-oblique mounting configuration, with the angular momentum direction of each individual reaction wheel perpendicular to the oblique mounting angle of the satellite's three axes. Step 2: Arrange a set of control moment gyroscopes for the satellite; determine the satellite axis that requires high maneuverability according to the mission requirements, and arrange the frame axis and 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 gyroscope, and the angular momentum directions of the two control moment gyroscope groups are kept in opposite states to avoid angular momentum interference caused by the control moment gyroscopes spinning in orbit; Step 3: Based on the hybrid control configuration of reaction wheel assembly and control moment gyroscope, realize the three-axis control of the satellite; in steady-state missions, use the three-wheel control strategy to achieve steady-state control of the satellite's three axes; in high-maneuver missions, use the control moment gyroscope assembly to achieve rapid maneuvering of a single axis. Step 4: When any reaction wheel malfunctions and causes the satellite to be unable to achieve three-axis steady-state control, switch the control logic, keep one of the control moment gyroscopes in a low-speed frame axis locked, and use the high-speed rotor part as a backup reaction wheel by changing its speed to maintain the satellite's three-axis steady-state control; when the control moment gyroscope is used as a backup reaction wheel, the satellite will only be controlled by the reaction wheel group.

2. The hybrid control configuration method for highly maneuverable micro / nano satellites according to claim 1, characterized in that: In step one, the installation matrix of the reaction wheel assembly is as follows:

3. The hybrid control configuration method for highly maneuverable micro / nano satellites according to claim 1, characterized in that: In step two, the installation matrix for the control torque gyroscope assembly is as follows: In this matrix, matrix A represents the direction of the angular momentum vector after the control moment gyroscope frame angle has rotated 90°, and matrix B represents the direction of the angular momentum vector when the control moment gyroscope frame angle is 0°.

4. The hybrid control configuration method for highly maneuverable micro / nano satellites according to claim 1, characterized in that: In step four, to ensure that the components of the high-speed control torque gyroscope's components along the celestial axis are equal, the direction of the high-speed rotor angular momentum of the control torque gyroscope is made to be at the same angle as the corresponding oblique angle of the celestial axis:

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