JP2009190506A - Posture control device for artificial satellite and posture control method of artificial satellite - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To solve a problem that a CMG (control moment gyro) which is an actuator for the posture control of an artificial satellite can output large torque but increases a torque resolution and a torque output error. <P>SOLUTION: This posture control device is provided with a CMG driving computing part 5 capable of outputting torque on an open loop and an RW driving control part to distribute torque to an RW (reaction wheel) 10 into the CMG in accordance with a gimbal angle of the CMG 4 and changes an open loop control mode by the CMG and a closed loop control mode by the RW over to each other at a maneuver time and at a non-maneuver time of the artificial satellite. <P>COPYRIGHT: (C)2009,JPO&INPIT

Description

  The present invention relates to an attitude control device for controlling the attitude of an artificial satellite using a control moment gyro, and an attitude control method for an artificial satellite.

  A control moment gyro (hereinafter referred to as CMG) is an actuator used for attitude control of an artificial satellite. Conventionally, actuators such as reaction wheels (hereinafter referred to as RW) and thrusters have been used for attitude control of artificial satellites, but CMG can generate a larger torque than these actuators. The mobility of the satellite is improved.

  The CMG can obtain an output torque by rotating a reaction wheel (hereinafter referred to as RW) that rotates at high speed around the gimbal axis. Therefore, the torque acting on the satellite from the CMG can be obtained by the outer product of the RW angular momentum vector and the gimbal angular velocity vector. At this time, the control method for generating a large torque by keeping the angular momentum of the RW constant and increasing the gimbal angular velocity. Is common.

  Also, in a control system using a plurality of CMGs, all output torque directions called singular points are included in one plane, and there is a state in which three-axis attitude control torque cannot be output. For this reason, for example, two CMGs and four RWs are provided in combination, and CMGs and RWs are independently controlled to construct a control system that is not affected by singularities (for example, Patent Document 1).

JP 2004-90796 A

  CMG has a torque resolution determined by the control performance of the gimbal angular velocity, and can output a large torque. On the other hand, compared with the torque output performance of RW used for attitude control of conventional satellites, CMG has a torque resolution and torque output error. There is a drawback that the control accuracy is reduced due to an increase in size.

  The attitude control device of Patent Document 1 aims to solve this problem by separately mounting both CMG and RW. To mount such a group of actuators, the space inside the artificial satellite Is limited and not realistic.

  In addition, although the RW installed in the CMG can output torque that is originally suitable for highly accurate attitude control, the method of use in the CMG is limited to the accumulation of angular momentum, and its capability is sufficient. It is not alive.

  Furthermore, in a control system using only CMG, a singularity problem occurs where torque cannot be output in a certain direction, and construction of a control system that avoids this singularity is extremely complicated and difficult.

  The present invention has been made in order to solve the above-mentioned problems, and can obtain the attitude control with high accuracy while utilizing the advantages of both CMG and RW while maintaining the high maneuverability of the satellite using CMG. With the goal.

  An attitude control device for a satellite according to the present invention includes a control moment gyro having a plurality of reaction wheels rotating around a reaction wheel rotation axis, and a plurality of gimbal shafts rotating the reaction wheels around a gimbal rotation axis. During maneuver of the satellite, the gimbal angle error and gimbal angular velocity error are calculated based on the target gimbal angle and target gimbal angular velocity command value, gimbal angle measured value and gimbal angular velocity measured value of the gimbal axis. The control moment gyro drive calculation unit that outputs the gimbal angle command value and the gimbal angular velocity command value for each reaction wheel and controls the angular momentum command value to each reaction wheel to be constant, and the gimbal when the satellite is not maneuvered. The gimbal angle of the axis is controlled to be constant, the attitude error is calculated based on the target attitude angle and attitude angle estimated value of the satellite, the target angular velocity and the estimated angular velocity value, and the calculated attitude error and the gimbal angle measurement of the gimbal axis are performed. A reaction wheel drive calculation unit that calculates a torque command value for each of the reaction wheels based on the value.

  According to this invention, the RW is mounted inside the CMG, and the attitude of the three axes is controlled by the RW when the artificial satellite is not maneuvered, and the MG is controlled by rotating the RW by the gimbal axis at the time of the maneuver. It is possible to change the attitude with high mobility without being affected, and to prevent the control accuracy of the satellite from being lowered.

Embodiment 1 FIG.
Embodiment 1 according to the present invention will be described below with reference to the drawings. FIG. 1 is a block diagram showing the configuration of the attitude control device for an artificial satellite according to the first embodiment.

  In the figure, the attitude control device includes an attitude maneuver planning unit 1, a regulator calculation unit 2, an RW drive calculation unit 3, a CMG 4, a CMG drive calculation unit 5, and an attitude determination unit 7. The attitude control device is mounted on the artificial satellite 6 and controls the attitude of the artificial satellite 6. The attitude maneuver planning unit 1 is a target gimbal angle / angular velocity command value composed of a target gimbal angle command value and a target gimbal angular velocity command value on the CMG gimbal axis for the artificial satellite 6 to obtain a target attitude according to the maneuver of the satellite. Calculate The maneuver plan of the artificial satellite is determined based on the operation plan of the artificial satellite by the operator of the ground station (not shown) that controls the artificial satellite, and the attitude maneuver planning unit 1 through the communication between the ground station and the artificial satellite. Is input. The regulator calculation unit 2 calculates a posture error from the target gimbal angle / angular velocity command value output from the posture maneuver planning unit 1 and the posture angle / angular velocity estimated value obtained from the posture determination unit 7. Calculate the desired attitude control torque. The regulator calculation unit 2 switches the calculation processing mode between maneuver and non-maneuver of the artificial satellite, and outputs different calculation results. The RW drive calculation unit 3 distributes the attitude control torque obtained from the regulator calculation unit 2 to each RW, and obtains a torque command value corresponding to each RW.

  FIG. 2 is a diagram showing a schematic configuration of the CMG 4. The CMG 4 includes an RW 10, a gimbal shaft 11 that rotates the RW 10 around the gimbal rotation axis, and a frame 12 that fixes the gimbal axis 11 to the satellite structure of the artificial satellite 6. The gimbal shaft 11 has a frame that rotates inside the frame 12. In the drawing, the frame that connects the RW 10 to the gimbal shaft 11 is indicated by a dotted line for convenience of illustration. The CMG 4 is an actuator that can drive the RW 10 around the gimbal axis 11, and outputs attitude variation torque to the artificial satellite by driving the CMG 4. The CMG 4 has an angle and angular velocity detector (not shown), and includes a gimbal angle measurement value indicating the angle of the gimbal shaft 11 with respect to the frame 12 and a gimbal angular velocity measurement value indicating the angular velocity of the gimbal shaft 11 with respect to the frame 12. Outputs gimbal angle and angular velocity measurement values. The CMG 4 includes a motor (not shown) that rotates the RW 10 around the gimbal rotation axis, a motor (not shown) that rotates the RW 10 around the gimbal axis 11, and a control circuit (not shown) that controls the rotation of the motor. However, the illustration and detailed description thereof are omitted here.

  The CMG drive calculation unit 5 calculates a target gimbal angle / angular velocity command value including a target gimbal angle command value and a target gimbal angular velocity command value for the gimbal axis, and a gimbal angle measurement value and a gimbal angular velocity measurement value for the gimbal axis obtained from the CMG 4. Based on the measured gimbal angle / angular velocity measurement value, a gimbal angle / angular velocity command value including a gimbal angle command value and a gimbal angular velocity command value for the gimbal axis is calculated. The attitude determination unit 7 calculates estimated values of the attitude angle and angular velocity of the artificial satellite based on sensor data such as an attitude angle detection gyro and an angular velocity detection gyro mounted on the artificial satellite 6.

Next, with respect to the attitude control device of the first embodiment, operation modes during maneuver and non-maneuver will be described.
First, the non-maneuver mode (RW drive mode) of the artificial satellite 6 will be described with reference to FIG. The CMG is not driven except when maneuvering, the gimbal angle is fixed, and three-axis drive control by RW is performed by closed loop control. At this time, the regulator calculation unit 2 calculates the three-axis attitude control torque of the artificial satellite 6 to compensate for the attitude error obtained from the target attitude angle and angular velocity of the artificial satellite 6 and the estimated attitude angle and angular velocity. Is done. The triaxial attitude control torque calculated by the regulator calculation unit 2 is distributed to the torque command value for each RW by the RW drive calculation unit 3.

FIG. 3 is a diagram showing an arrangement example of CMGs 4 and RWs installed in the CMG, and FIG. 4 shows each RW with respect to the triaxial fixed coordinate system (X, Y, Z) of the artificial satellite 6 in the CMG arrangement of FIG. It is a figure which shows the example of arrangement | positioning of a rotating shaft.
FIG. 3 shows an example of the pyramid type arrangement of four RWs mounted inside the CMG using four CMGs. In this case, the rotational axis of the RW is open with respect to the Z axis of the artificial satellite 6, and the initial position of the gimbal angle is set so that the respective rotational axes are not parallel. As a result, by using a combination of RWs mounted in each CMG, it is possible to output torque that can be attitude controlled in all three axes of the artificial satellite 6.
Since the gimbal angle changes during maneuver of the artificial satellite 6, the rotation axis of the RW also changes, but after the maneuver, the gimbal angle almost returns to the initial position. For this reason, three-axis attitude control by RW is always possible except during maneuver.

  Here, the relationship between the three-axis attitude control torque (τx, τy, τz) of the artificial satellite 6 and the torque (τwi, i: number of RWs) of each RW is generally expressed by the following equation. .

  A is a 3 × i torque distribution matrix uniquely determined geometrically by the arrangement of RWs. If the gimbal angle of the gimbal axis 11 is φj (j: number of CMGs), the arrangement of RW changes according to φj of each CMG, so each element of the matrix A is a function of φj as follows.

  The gimbal angle measurement value 8 in FIG. 1 corresponds to φj, and this value is fed back to the RW drive calculation unit 3, so that a matrix A that changes after maneuver can be obtained. Using the obtained matrix A, the distribution torque to each RW can be obtained from the attitude control torque of the artificial satellite 6.

Thus, in the non-maneuver mode, the three-axis attitude control of the artificial satellite 6 can be performed using only the RW, and the attitude control with the torque output accuracy of the RW can be performed.

Next, the maneuver mode (CMG drive mode) of the artificial satellite 6 will be described.
When maneuver of the artificial satellite 6 is performed, the RW angular momentum command value is constant, the RW is controlled at a constant rotation, and the gimbal angle of the CMG 4 is controlled according to the target gimbal angle and angular velocity obtained by the attitude maneuver plan 1 of FIG. To do. At this time, as shown in FIG. 1, the open loop control of the CMG 4 is performed by cutting the closed loop that compensates for the posture error.

  Since CMG4 is open-loop control, it is possible to perform trajectory planning without considering singular points when planning target gimbal angles and angular velocities, and it is necessary to construct a complex control system required when passing singular points Not.

  In addition, since CMG4 performs maneuver with open loop control, an attitude error occurs immediately after maneuver, but as described above, after switching to three-axis attitude control by RW having a closed loop that compensates for attitude error, as described above, It is possible to remove the posture error.

  As described above, the attitude control device for a satellite according to this embodiment includes a plurality of reaction wheels that rotate around the reaction wheel rotation axis, and a plurality of gimbal shafts that rotate the reaction wheels around the gimbal rotation axis, respectively. The gimbal angle error and gimbal angular velocity error are calculated based on the target gimbal angle, target gimbal angular velocity command value, gimbal angle measured value, and gimbal angular velocity measured value of the gimbal axis when maneuvering the control moment gyro with A control moment gyro drive arithmetic unit that outputs a gimbal angle command value and a gimbal angular velocity command value to each gimbal rotation axis, and controls the angular momentum command value to each reaction wheel to be constant, and an artificial satellite of During maneuver, the gimbal angle of the gimbal axis is controlled to be constant, and the attitude error is calculated based on the target attitude angle and attitude angle estimated value of the artificial satellite and the target angular velocity and angular velocity estimated value, and the calculated attitude error and the gimbal are calculated. A reaction wheel drive calculation unit that calculates a torque command value for each reaction wheel based on a measured value of the shaft gimbal angle.

  With this configuration, not only CMG4 but also the RW installed in CMG can be fully utilized, and high-precision attitude control can be achieved with relatively simple logic while maintaining the high maneuverability of the satellite. Can be realized.

  In addition, the RW is installed in the CMG, and the RW drive mode in which the attitude of the three axes is controlled by the RW when the artificial satellite is not maneuvered, and the CMG drive mode in which the MG is rotated by the gimbal axis to control the CMG4 in the maneuver. By switching the two modes and controlling the attitude of the artificial satellite, it is possible to prevent the influence of the singular point and to prevent the control accuracy of the artificial satellite from being lowered.

  Further, by executing a mode (CMG drive mode) in which CMG 4 is open-loop controlled at the time of maneuver, a control plan can be made without considering singular points. Further, since the mode is switched to the RW three-axis attitude control mode (RW drive mode) having a closed loop that compensates for the attitude error after the maneuver, it is possible to remove the attitude error generated during the maneuver by the open loop control.

  Also, by using the RW inside the CMG while taking advantage of both the CMG and RW, the number of CMGs and RWs combined is reduced to the minimum (at least 4 CMGs if redundancy is not considered). It is possible.

It is a block diagram of the attitude | position control apparatus of the artificial satellite in Embodiment 1 of this invention. It is the schematic of CMG in Embodiment 1 of this invention. It is an example of the layout of CMG and CMG internal mounting RW in Embodiment 1 of this invention. The RW rotation axis with respect to the triaxial fixed coordinate system (X, Y, Z) of the satellite corresponding to the CMG arrangement in the first embodiment of the present invention is shown.

Explanation of symbols

1 attitude maneuver planning unit, 2 regulator calculation unit, 3 RW drive calculation unit, 4 CMG, 5 CMG drive calculation unit, 6 artificial satellite, 7 attitude determination unit, 10 RW, 11 gimbal axis, 12 frames.

Claims (2)

A control moment gyro having a plurality of reaction wheels that rotate around a reaction wheel rotation axis, and a plurality of gimbal axes that rotate the reaction wheels around a gimbal rotation axis, respectively.
At the time of maneuver of the artificial satellite, the gimbal angle error and the gimbal angular velocity error are calculated based on the target gimbal angle and the target gimbal angular velocity command value, the gimbal angle measured value and the gimbal angular velocity measured value of the gimbal axis. A control moment gyro drive calculation unit that outputs a gimbal angle command value and a gimbal angular velocity command value and controls the angular momentum command value to each reaction wheel to be constant,
The gimbal angle of the gimbal axis is controlled to be constant when the satellite is not maneuvered, and the attitude error is calculated based on the target attitude angle and attitude angle estimated value of the satellite and the target angular velocity and angular velocity estimated value. A reaction wheel drive calculation unit that calculates a torque command value to each reaction wheel based on the error and the gimbal angle measurement value of the gimbal shaft;
An attitude control device for artificial satellites. Using multiple control moment gyros that rotate the reaction wheel provided inside around the gimbal rotation axis,
During maneuver of the satellite, the gimbal angle error and the gimbal angular velocity error are calculated based on the target gimbal angle of the gimbal axis, the target gimbal angular velocity command value, the gimbal angle measured value, and the gimbal angular velocity measured value. A control moment gyro drive mode for outputting an angle command value and a gimbal angular velocity command value and controlling the angular momentum command value to each reaction wheel to be constant;
The gimbal angle of the gimbal axis is controlled to be constant when the satellite is not maneuvered, and the attitude error is calculated based on the target attitude angle and attitude angle estimated value and target angular velocity and angular velocity estimated value of the satellite. Further, the attitude control of the artificial satellite is performed by switching between two modes, a reaction wheel driving mode for calculating a torque command value to each reaction wheel based on the measured gimbal angle value of the gimbal axis. Artificial satellite attitude control method.

Images