JP2018027761A - Attitude control device, space craft, attitude control method and attitude control program - Google Patents

Abstract

PROBLEM TO BE SOLVED: To control attitudes simultaneously in three axes using a small number of control moment gyros.SOLUTION: Two control moment gyros 13 of an attitude control device 10 individually have a wheel 31 and a gimbals 32. A gyro calculation unit 24 of the attitude control device 10 calculates a command value for each of the control moment gyros 13 as for an angular speed of the gimbals from a set value of a torque in a three-axes direction for controlling the attitude. The gyro calculation unit 24 calculates a common change amount for an angular momentum of a wheel from a set value of a torque in a three-axes direction for controlling the attitude, then weighs the calculated common change amount with a coefficient for each of the control moment gyros 13, and calculates a command value for each of the control moment gyros 13.SELECTED DRAWING: Figure 1

Description

  The present invention relates to an attitude control device, a spacecraft, an attitude control method, and an attitude control program.

  Patent Document 1 discloses an attitude control device that performs three-axis attitude control of an artificial satellite using two CMGs. “CMG” is an abbreviation for control moment gyro.

  Non-Patent Document 1 discloses a three-axis attitude control method for an artificial satellite using VSCMG. “VSCMMG” is an abbreviation for Variable Speed CMG.

JP 2009-173069 A

Hansetter Schaub, Srinivas R. Vadali and John L. Junkins, “Feedback Control Law for Variable Speed Control Moments Gyros”, Journal of the Agricultural Sciences, Vol. 46, no. 3, July-Sept. , 1998, Pages 307-328.

  In the attitude control of satellites using CMG, the CMG wheel generally only accumulates angular momentum, and the gyro torque generated when the wheel angular momentum direction is changed by controlling the rotation angle of the gimbal shaft. Is used. When used in this way, the degree of freedom of control of the CMG is one, and in order to realize the three-axis attitude control of the artificial satellite, it is necessary to configure three or more units. Furthermore, considering the redundant system, it is desirable to have a configuration of four or more CMGs. Therefore, the increase in the number of installed CMGs prevents the miniaturization and weight reduction of the satellite.

  In the system disclosed in Patent Document 1, only the torque output by controlling the rotation angle of the gimbal shaft by rotating the CMG wheel at a constant speed is used. As described above, in the attitude control device of Patent Document 1, attitude control is performed using the fact that the output torque direction changes due to the rotation of the gimbal shaft, but the torque direction that two CMGs can output instantaneously is Since it is fixed on a plane, the three-axis attitude of the artificial satellite cannot be controlled simultaneously.

  In the method disclosed in Non-Patent Document 1, in addition to the torque output by controlling the rotation angle of the gimbal shaft, the torque generated by changing the angular momentum of the wheel is also used, and thereby the artificial satellite is used. The three-axis attitude control torque is output. In VSCMG as in Non-Patent Document 1, torque is output when the wheel angular momentum is changed. Therefore, the rotational angle of the gimbal shaft and the wheel angular momentum can be controlled, and the control freedom of CMG is 2 per unit. Become. Therefore, the degree of freedom of control becomes 2 × 2 = 4 with two CMGs, and the three-axis attitude control of the artificial satellite becomes possible. Since there are three control axes for four degrees of freedom of control, one degree of freedom is redundant, so the three-axis attitude control torque is not uniquely distributed to the two CMGs as it is, and appropriate control is performed. May not be done. In addition, the calculation processing for controlling both the rotation angle of the gimbal shaft and the wheel angular momentum becomes a heavy load, and the processing time may be increased. When the processing time becomes long, it becomes impossible to output a command command to the CMG at an appropriate timing, and the responsiveness and stability of the attitude control system deteriorate.

  An object of the present invention is to perform three-axis simultaneous attitude control by light load processing using a small number of control moment gyros.

An attitude control device according to one aspect of the present invention includes:
At least two control moment gyros individually having a rotating wheel and a gimbal that rotates around a rotation axis that is different in direction from the rotation axis of the wheel while holding the wheel;
From the set values of torque in the three-axis directions for attitude control, the command value for each control moment gyro is calculated for the angular velocity of the gimbal, and the common change amount is calculated for the angular momentum of the wheel. And a gyro operation unit that calculates a command value for each control moment gyro by weighting the calculated common change amount with a coefficient for each control moment gyro.

  In the present invention, the command value for each control moment gyro is calculated for the angular velocity of the gimbal from the set value of the torque in the three-axis directions for posture control, and the common amount of change is calculated for the angular momentum of the wheel. After that, the common change amount is weighted by a coefficient for each control moment gyro to calculate a command value for each control moment gyro. For this reason, it is possible to perform the attitude control at the same time for the three axes by light load processing using a small number of control moment gyros.

FIG. 2 is a block diagram illustrating a configuration of the attitude control device according to the first embodiment. FIG. 3 shows a configuration of a control moment gyro according to the first embodiment. FIG. 3 is a diagram illustrating an arrangement example of two control moment gyros according to the first embodiment. FIG. 3 is a diagram illustrating an arrangement example of two control moment gyros according to the first embodiment. 5 is a flowchart showing the operation of the attitude control device according to the first embodiment. FIG. 6 is a block diagram showing a configuration of an attitude control device according to a modification of the first embodiment.

  Hereinafter, embodiments of the present invention will be described with reference to the drawings. In addition, the same code | symbol is attached | subjected to the part which is the same or it corresponds in each figure. In the description of the embodiments, the description of the same or corresponding parts will be omitted or simplified as appropriate.

Embodiment 1 FIG.
This embodiment will be described with reference to FIGS.

*** Explanation of configuration ***
With reference to FIG. 1, the structure of the attitude | position control apparatus 10 which concerns on this Embodiment is demonstrated.

  The attitude control device 10 is a device that is mounted on the spacecraft 40 and controls the attitude of the spacecraft 40. The spacecraft 40 is an artificial satellite in the present embodiment, but may be another type of spacecraft such as a space station.

  In the present embodiment, posture control apparatus 10 is a computer. The attitude control device 10 includes a processor 11 and other hardware such as a memory 12 and a control moment gyro 13. The processor 11 is connected to other hardware via a signal line, and controls these other hardware.

  The attitude control device 10 includes an attitude estimation unit 21, an attitude calculation unit 22, a torque calculation unit 23, and a gyro calculation unit 24 as functional elements. Functions of “units” such as the posture estimating unit 21, the posture calculating unit 22, the torque calculating unit 23, and the gyro calculating unit 24 are realized by software.

  The processor 11 is an IC that performs processing. “IC” is an abbreviation for Integrated Circuit. Specifically, the processor 11 is a CPU. “CPU” is an abbreviation for Central Processing Unit.

  Specifically, the memory 12 is a flash memory or a RAM. “RAM” is an abbreviation for Random Access Memory.

  The control moment gyro 13 will be described later.

  The attitude control device 10 may include a communication device as hardware.

  The communication device includes a receiver that receives data and a transmitter that transmits data. Specifically, the communication device is a communication chip or a NIC. “NIC” is an abbreviation for Network Interface Card.

  The memory 12 stores a program for realizing the function of “unit”. This program is read into the processor 11 and executed by the processor 11. The memory 12 also stores an OS. “OS” is an abbreviation for Operating System. The processor 11 executes a program that realizes the function of “unit” while executing the OS. A part or all of the program for realizing the function of “unit” may be incorporated in the OS.

  The program and OS that realize the function of “unit” may be stored in the auxiliary storage device. Specifically, the auxiliary storage device is a flash memory or an HDD. “HDD” is an abbreviation for Hard Disk Drive. The program and OS stored in the auxiliary storage device are loaded into the memory 12 and executed by the processor 11.

  The attitude control device 10 may include a plurality of processors that replace the processor 11. The plurality of processors share the execution of a program that realizes the function of “unit”. Each processor is an IC that performs processing in the same manner as the processor 11.

  Information, data, signal values, and variable values indicating the result of the processing of “part” are stored in the memory 12, the auxiliary storage device, or a register or cache memory in the processor 11.

  A program for realizing the function of “unit” may be stored in a portable recording medium such as a magnetic disk or an optical disk.

  The attitude control device 10 includes two control moment gyros 13 in the present embodiment, but the number of control moment gyros 13 is not limited to two and may be three or more. That is, the attitude control device 10 only needs to include at least two control moment gyros 13.

  The configuration of each control moment gyro 13 will be described with reference to FIG.

  The control moment gyro 13 has a wheel 31 and a gimbal 32 individually.

  The wheel 31 rotates around the wheel rotation shaft 33. The wheel 31 is also called a rotor.

  The gimbal 32 rotates around the gimbal rotation shaft 34 while holding the wheel 31. The gimbal rotating shaft 34 is a rotating shaft having a direction different from that of the wheel rotating shaft 33 that is the rotating shaft of the wheel 31. The gimbal 32 includes a frame 35 fixed to the structure of an artificial satellite that is a spacecraft 40, a shaft 36 that rotates around the gimbal rotation axis 34 with respect to the frame 35, and a motor (not shown) that drives the shaft 36. And have. The shaft 36 is connected to the wheel 31 inside the frame 35.

  The control moment gyro 13 outputs torque by rotating the wheel 31 rotating around the wheel rotation axis 33 around the gimbal rotation axis 34 with respect to the frame 35 fixed to the structure of the artificial satellite. Further, the control moment gyro 13 outputs torque in the direction of the wheel rotation shaft 33 by changing the rotation angular velocity of the wheel 31.

  Although not shown, the control moment gyro 13 includes an angle detector, and outputs a measurement value of a gimbal angle that is a rotation angle of the shaft 36 with respect to the frame 35 using the angle detector. Further, the control moment gyro 13 has an angular velocity detector, and outputs a measured value of the wheel angular momentum obtained from the measured value of the rotational angular velocity of the wheel 31 around the wheel rotation axis 33 using the angular velocity detector. However, since the measured value of the wheel angular momentum can be replaced by the command value of the wheel angular momentum calculated by the gyro calculating unit 24 described later, the angular velocity detector is not essential.

  With reference to FIG. 3 and FIG. 4, the example of arrangement | positioning of the two control moment gyros 13 applied to this Embodiment is demonstrated. Here, one control moment gyro 13 is referred to as CMG # 1, and the other control moment gyro 13 is referred to as CMG # 2.

In this example, CMG # 1 and CMG # 2 have the gimbal rotation shaft 34 and the wheel rotation shaft 33 orthogonal to each other. When the gimbal angle of CMG # 1 is 0 °, the direction of the wheel rotation axis 33 of CMG # 1 coincides with the + X-axis direction of the spacecraft 40. When the gimbal angle of CMG # 2 is 0 °, the direction of the wheel rotation axis 33 of CMG # 2 coincides with the −X axis direction of the spacecraft 40. CMG direction # 1 of gimbal pivot axis 34 coincides with the + Y-axis of the spacecraft 40 in a direction rotated by alpha ₁ about the + X-axis of the spacecraft 40. CMG direction # 2 of gimbal pivot axis 34 coincides with the + Y-axis of the spacecraft 40 in a direction rotated by alpha ₂ around the + X-axis of the spacecraft 40.

*** Explanation of operation ***
With reference to FIG. 5, the operation of the attitude control apparatus 10 according to the present embodiment will be described. The operation of the attitude control device 10 corresponds to the attitude control method according to the present embodiment. The operation of the attitude control device 10 corresponds to the processing procedure of the attitude control program according to the present embodiment.

  In step S <b> 1, the posture estimation unit 21 estimates the posture of the spacecraft 40.

  Specifically, the attitude estimation unit 21 calculates estimated values of the attitude angle and attitude angular velocity of the spacecraft 40 from data of sensors such as an attitude quaternion detection sensor and an angular velocity detection gyro mounted on the spacecraft 40. Then, the posture estimation unit 21 outputs the calculated estimated value to the register or the memory 12 of the processor 11.

  In step S <b> 2, the attitude calculation unit 22 calculates the target attitude of the spacecraft 40.

  Specifically, the attitude calculation unit 22 is set in advance, or in accordance with a destination commanded from outside, a target value of the attitude angle and attitude angular velocity of the spacecraft 40 and a set value of the feedforward torque Calculate Then, the posture calculation unit 22 outputs the calculated target value and set value to the register or the memory 12 of the processor 11.

  In step S <b> 3, the torque calculation unit 23 calculates a set value of torque in the three-axis directions for attitude control of the spacecraft 40 from the difference between the attitude estimated by the attitude estimation unit 21 and the target attitude.

  Specifically, the torque calculator 23 subtracts the estimated values of the posture angle and posture angular velocity output in step S1 from the target values of the posture angle and posture angular velocity output in step S2, respectively, to obtain the posture angle and posture angle. Find the angular velocity error. The torque calculation unit 23 calculates a set value of the attitude control torque necessary for the attitude control of the spacecraft 40 from the obtained error and the set value of the feedforward torque output in step S2. Then, the torque calculator 23 outputs the calculated set value to the register of the processor 11 or the memory 12.

  In step S <b> 4, the gyro calculating unit 24 calculates a command value for each control moment gyro 13 for the angular velocity of the gimbal 32 from the set value of the torque in the three axial directions for posture control. Further, the gyro calculating unit 24 calculates a common change amount for the angular momentum of the wheel 31 from the set values of torques in the three axial directions for posture control, and then calculates the calculated common change amount as a control moment gyro. A command value for each control moment gyro 13 is calculated by weighting with a coefficient for each 13. In the present embodiment, at this time, the gyro operation unit 24 calculates the sum of the common reference value and the value obtained by weighting the calculated common change amount by the coefficient for each control moment gyro 13 for the angular momentum of the wheel 31. The control moment gyro 13 is calculated as a command value.

  Specifically, the gyro operation unit 24 calculates CMG # 1 based on the setting value output in step S3 and the measured value of the gimbal angle obtained from CMG # 1 and CMG # 2 by a calculation method described later. The command value of the gimbal angular velocity, the command value of the gimbal angular velocity of CMG # 2, the command value of the wheel torque of CMG # 1, and the command value of the wheel torque of CMG # 2. Then, the gyro operation unit 24 outputs the calculated command value to the register of the processor 11 or the memory 12.

  The gyro operation unit 24 may calculate a command value of wheel angular momentum obtained by integrating the command value of wheel torque instead of the command value of wheel torque. In the present embodiment, calculating the command value of the wheel torque, or calculating the command value of the wheel angular momentum corresponds to calculating the command value for each control moment gyro 13 with respect to the angular momentum of the wheel 31.

  Further, as described above, since the measured value of the wheel angular momentum can be replaced by the command value of the wheel angular momentum, the gyro calculating unit 24 uses the command value of the wheel angular momentum in the present embodiment. Although the command value of the gimbal angular velocity is calculated, the command value of the gimbal angular velocity may be calculated using the measured value of the wheel angular momentum. That is, the gyro operation unit 24 measures the setting value output in step S3, the measured value of the gimbal angle obtained from CMG # 1 and CMG # 2, and the measurement of the wheel angular momentum obtained from CMG # 1 and CMG # 2. CMG # 1 gimbal angular velocity command value, CMG # 2 gimbal angular velocity command value, CMG # 1 wheel torque command value, and CMG # 2 wheel torque command value. You may calculate.

  In step S5, the gyro calculating unit 24 rotates the wheel 31 and the gimbal 32 of each control moment gyro 13 according to the command value for each control moment gyro 13 calculated with respect to the angular velocity of the gimbal 32 and the angular momentum of the wheel 31. Let Each control moment gyro 13 outputs torque in the three-axis directions for attitude control to the spacecraft 40 by the rotation of the wheel 31 and the gimbal 32.

  Specifically, the gyro calculating unit 24 transmits the CMG # 1 gimbal angular velocity command value and the CMG # 1 wheel torque command value output in step S4 to the CMG # 1. CMG # 1 receives the transmitted command value, and rotates wheel 31 and gimbal 32 of CMG # 1 according to the received command value. Similarly, the gyro operation unit 24 transmits the command value for the CMG # 2 gimbal angular velocity and the command value for the wheel torque of CMG # 2 output in step S4 to the CMG # 2. CMG # 2 receives the transmitted command value, and rotates wheel 31 and gimbal 32 of CMG # 2 in accordance with the received command value.

  Below, the specific calculation method of step S4 is demonstrated.

The triaxial attitude control torque output by CMG # 1 and CMG # 2 includes the CMG # 1 gimbal angular velocity, the CMG # 2 gimbal angular velocity, the time change rate of the CMG # 1 wheel angular momentum, and the CMG # 2. And the time rate of change of the wheel angular momentum of The “time rate of change of wheel angular momentum” is synonymous with wheel torque.

Here, A is a 3 × 4 matrix determined by the arrangement of the control moment gyros 13 and is a function of the gimbal angle and the wheel angular momentum of each control moment gyro 13. In order to calculate a command value to the control moment gyro 13 from the three-axis attitude control torque, it is necessary to calculate an inverse matrix of A. Since there is no inverse matrix of a non-regular matrix such as a 3 × 4 matrix, a command value to the control moment gyro 13 can be calculated by using a pseudo inverse matrix such as Equation 2.

However, in a computer such as an FPGA or embedded software with a limited processing capacity, the calculation of the pseudo inverse matrix becomes a heavy load and the processing time becomes long. “FPGA” is an abbreviation for Field-Programmable Gate Array. If the processing time becomes long, the command value to the control moment gyro 13 cannot be output at an appropriate timing, and the responsiveness and stability of the attitude control system deteriorate. Therefore, in the present embodiment, by providing a constant relational expression for the wheel angular momentum of each control moment gyro 13, the triaxial attitude control torque is distributed to the two control moment gyros 13. Specifically, the reference value of the wheel angular momentum is h _c , the common change amount of the wheel angular momentum is Δh, the wheel angular momentum of CMG # 1 is h ₁ = h _c + k ₁ Δh, and the wheel angular momentum of CMG # 2. Is set to h ₂ = h _c + k ₂ Δh. h _c is a fixed value. In this case, it is possible to provide a CMG # 1 and CMG # 2 of the wheel angular momentum by giving different arbitrary values in coefficient parameters k ₁ and k ₂ the relationship. By giving this relational expression, the three-axis attitude control torque of the spacecraft 40 can be expressed by Expression 3.

  Here, B is a 3 × 3 matrix determined by the arrangement of the control moment gyros 13 and is a function of the gimbal angle and the wheel angular momentum of each control moment gyro 13. Since the time differential value of the wheel angular momentum expresses the torque, the wheel torque can be calculated based on the time differential value of Δh. Since the inverse of the 3 × 3 matrix can be solved more easily than the calculation of the pseudo inverse of the 3 × 4 matrix, Equation 3 can be expressed as the gimbal angular velocity of CMG # 1, the gimbal angular velocity of CMG # 2, The result solved for the time derivative of Δh is a simple mathematical formula. The gyro operation unit 24 processes mathematical formulas obtained by solving Equation 3 for the CMG # 1 gimbal angular velocity, the CMG # 2 gimbal angular velocity, and the time differential value of Δh, thereby obtaining each of the desired attitude control torques. A command value for the gimbal angular velocity of the control moment gyro 13 and a command value for the wheel torque are obtained.

As described above, in this embodiment, when the gimbal angle of CMG # 1 is 0 °, the direction of the wheel rotation axis 33 of CMG # 1 coincides with the + X-axis direction of the spacecraft 40. When the gimbal angle of CMG # 2 is 0 °, the direction of the wheel rotation axis 33 of CMG # 2 coincides with the −X axis direction of the spacecraft 40. CMG direction # 1 of gimbal pivot axis 34 coincides with the + Y-axis of the spacecraft 40 in a direction rotated by alpha ₁ about the + X-axis of the spacecraft 40. CMG direction # 2 of gimbal pivot axis 34 coincides with the + Y-axis of the spacecraft 40 in a direction rotated by alpha ₂ around the + X-axis of the spacecraft 40. At this time, A in Expression 3 is expressed by Expression 4.

In Equation 4, when θ ₁ = 0 and θ ₂ = 0, the pseudo inverse matrix of A is expressed as Equation 5.

At this time, the command value to the control moment gyro 13 is calculated as shown in Equation 6.

On the other hand, with the same control moment gyro 13 arrangement, B in Equation 3 is expressed as Equation 7.

By calculating the inverse matrix of B in Expression 7, the command value to the control moment gyro 13 is expressed as Expression 8.

In Equation 8, if θ ₁ = 0, θ ₂ = 0, k ₁ = −1, and k ₂ = 1, the command value to the control moment gyro 13 is calculated as in Equation 9.

Using the time differential value of Δh obtained in Equation 9, the time change rate of the wheel angular momentum of CMG # 1 and the time change rate of the wheel angular momentum of CMG # 2 are calculated as shown in Equation 10.

  Comparing Equation 6 with Equation 9 and Equation 10, the calculation results are exactly the same. That is, Expression 8 simplifies calculation of the command value to the control moment gyro 13 by the pseudo inverse matrix when both the CMG # 1 and CMG # 2 gimbal angles are close to 0 °. The gyro operation unit 24 can calculate the gimbal angular velocity command value and the wheel torque command value of each control moment gyro 13 from the desired attitude control torque by calculating Expression 8.

  With such a calculation method, it is possible to perform the three-axis attitude control of the spacecraft 40 by the two control moment gyros 13 by a simple calculation without solving the pseudo inverse matrix of the 3 × 4 matrix. This makes it possible to simultaneously control the three-axis attitudes of the spacecraft 40 such as an artificial satellite with light load processing while minimizing the number of control moment gyros 13 mounted. In addition, by providing a constant relational expression for the angular momentum of each control moment gyro 13 for three or more control moment gyros 13, the order of the matrix to be solved can be reduced, so the processing load is lightened. can do.

*** Explanation of the effect of the embodiment ***
In the present embodiment, the command value for each control moment gyro 13 is calculated for the angular velocity of the gimbal 32 from the set values of the torques in the three axial directions for posture control, and the angular momentum of the wheel 31 is common. After the change amount is calculated, the command value for each control moment gyro 13 is calculated by weighting the common change amount by the coefficient for each control moment gyro 13. For this reason, using a small number of control moment gyros 13, three-axis simultaneous attitude control can be performed by light load processing.

  In other words, in the present embodiment, the wheel angular momentum of each control moment gyro 13 is made variable and a certain relational expression is provided so that the three-axis attitude control torque is applied to at least two control moment gyros 13. Distributed. Specifically, for at least two control moment gyros 13, each wheel angular momentum is a value obtained by adding a reference value of the wheel angular momentum and a value obtained by multiplying the common change amount by a determined coefficient. By providing the relational expression as described above, the gimbal angular velocity command and the wheel torque command or the wheel angular momentum command for the control moment gyro 13 are calculated from the attitude control torque of the artificial satellite. For this reason, the attitude control of the three axes of the spacecraft 40 such as an artificial satellite with the minimum number of control moment gyros 13 mounted can be performed by light load processing. The minimum number of units is two if redundancy is not taken into consideration.

*** Other configurations ***
In the present embodiment, the function of “unit” is realized by software, but as a modification, the function of “unit” may be realized by hardware. About this modification, the difference with this Embodiment is mainly demonstrated.

  With reference to FIG. 6, the structure of the attitude | position control apparatus 10 which concerns on the modification of this Embodiment is demonstrated.

  The attitude control device 10 includes hardware such as a processing circuit 14 and a control moment gyro 13.

  The processing circuit 14 is a dedicated electronic circuit that realizes the function of “unit”. Specifically, the processing circuit 14 is a single circuit, a composite circuit, a programmed processor, a processor programmed in parallel, a logic IC, a GA, an FPGA, or an ASIC. “GA” is an abbreviation for Gate Array. “ASIC” is an abbreviation for Application Specific Integrated Circuit.

  The attitude control device 10 may include a plurality of processing circuits that replace the processing circuit 14. As a whole, the function of “unit” is realized by the plurality of processing circuits. Each processing circuit is a dedicated electronic circuit like the processing circuit 14.

  As another modification, the function of “unit” may be realized by a combination of software and hardware. That is, a part of the function of “unit” may be realized by a dedicated electronic circuit, and the rest may be realized by software.

  The processor 11, the memory 12, and the processing circuit 14 are collectively referred to as a “processing circuit”. That is, regardless of the configuration of the attitude control device 10 shown in FIGS. 1 and 6, the function of “unit” is realized by a processing circuit.

  “Part” may be read as “process”, “procedure”, or “processing”.

  As mentioned above, although embodiment of this invention was described, you may implement this embodiment partially. Specifically, only some of the functional elements of the posture control apparatus 10 according to this embodiment may be employed. In addition, this invention is not limited to this embodiment, A various change is possible as needed.

  DESCRIPTION OF SYMBOLS 10 Attitude control apparatus, 11 Processor, 12 Memory, 13 Control moment gyro, 14 Processing circuit, 21 Attitude estimation part, 22 Attitude calculation part, 23 Torque calculation part, 24 Gyro calculation part, 31 Wheel, 32 Gimbal, 33 Wheel rotation axis , 34 Gimbal rotation axis, 35 frame, 36 shaft, 40 spacecraft.

Claims (6)

At least two control moment gyros individually having a rotating wheel and a gimbal that rotates around a rotation axis that is different in direction from the rotation axis of the wheel while holding the wheel;
From the set values of torque in the three-axis directions for attitude control, the command value for each control moment gyro is calculated for the angular velocity of the gimbal, and the common change amount is calculated for the angular momentum of the wheel. An attitude control device comprising a gyro operation unit that calculates a command value for each control moment gyro by weighting the calculated common change amount with a coefficient for each control moment gyro.   For the angular momentum of the wheel, the gyro operation unit calculates a sum of a common reference value and a value obtained by weighting the calculated common change amount by a coefficient for each control moment gyro as a command value for each control moment gyro. The attitude control device according to claim 1. An attitude estimation unit for estimating the attitude of the spacecraft;
The torque calculation part which calculates the setting value of the torque of the said 3 axis direction for the attitude | position control of the said spacecraft from the difference between the attitude | position estimated by the said attitude | position estimation part and a target attitude | position is further provided. 2. The attitude control device according to 2.   A spacecraft on which the attitude control device according to any one of claims 1 to 3 is mounted. In an attitude control method using at least two control moment gyros each having a rotating wheel and a gimbal that rotates around a rotation axis that is different from the rotation axis of the wheel while holding the wheel,
From the set values of torque in the three-axis directions for attitude control, the command value for each control moment gyro is calculated for the angular velocity of the gimbal, and the common change amount is calculated for the angular momentum of the wheel. Calculate the command value for each control moment gyro by weighting the calculated common change amount with the coefficient for each control moment gyro.
A posture control method for rotating the wheel and the gimbal according to a command value for each control moment gyro calculated with respect to the angular velocity of the gimbal and the angular momentum of the wheel. A computer that calculates command values for at least two control moment gyros each having a rotating wheel and a gimbal that rotates around a rotation axis that is different from the rotation axis of the wheel while holding the wheel. In addition,
From the set values of torque in the three-axis directions for attitude control, the command value for each control moment gyro is calculated for the angular velocity of the gimbal, and the common change amount is calculated for the angular momentum of the wheel. An attitude control program that executes a process of calculating a command value for each control moment gyro by weighting the calculated common change amount with a coefficient for each control moment gyro.

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