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US20220388692A1 - Attitude determination and control system based on a quaternion kalman filter and method thereof - Google Patents

Attitude determination and control system based on a quaternion kalman filter and method thereof Download PDF

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US20220388692A1
US20220388692A1 US17/464,043 US202117464043A US2022388692A1 US 20220388692 A1 US20220388692 A1 US 20220388692A1 US 202117464043 A US202117464043 A US 202117464043A US 2022388692 A1 US2022388692 A1 US 2022388692A1
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attitude
satellite
actuator
torque
inertia
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Po-Hsun Yen
Shang-Jung Lee
Yung-Cheng Chang
Sung-Liang Hou
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Tensor Tech Co Ltd
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Tensor Tech Co Ltd
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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B13/00—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion
    • G05B13/02—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric
    • G05B13/04—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/244—Spacecraft control systems
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/244—Spacecraft control systems
    • B64G1/245—Attitude control algorithms for spacecraft attitude control
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B19/00—Programme-control systems
    • G05B19/02—Programme-control systems electric
    • G05B19/18—Numerical control [NC], i.e. automatically operating machines, in particular machine tools, e.g. in a manufacturing environment, so as to execute positioning, movement or co-ordinated operations by means of programme data in numerical form
    • G05B19/4155—Numerical control [NC], i.e. automatically operating machines, in particular machine tools, e.g. in a manufacturing environment, so as to execute positioning, movement or co-ordinated operations by means of programme data in numerical form characterised by programme execution, i.e. part programme or machine function execution, e.g. selection of a programme
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/26—Guiding or controlling apparatus, e.g. for attitude control using jets
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/26—Guiding or controlling apparatus, e.g. for attitude control using jets
    • B64G1/262—Guiding or controlling apparatus, e.g. for attitude control using jets having adjustable angles, e.g. gimbaled thrusters
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/28—Guiding or controlling apparatus, e.g. for attitude control using inertia or gyro effect
    • B64G1/283—Guiding or controlling apparatus, e.g. for attitude control using inertia or gyro effect using reaction wheels
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/28—Guiding or controlling apparatus, e.g. for attitude control using inertia or gyro effect
    • B64G1/286—Guiding or controlling apparatus, e.g. for attitude control using inertia or gyro effect using control momentum gyroscopes (CMGs)
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/32—Guiding or controlling apparatus, e.g. for attitude control using earth's magnetic field
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
    • B64G1/361—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using star sensors
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
    • B64G1/366—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using magnetometers
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/38—Guiding or controlling apparatus, e.g. for attitude control damping of oscillations, e.g. nutation dampers
    • B64G2001/245—
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B2219/00—Program-control systems
    • G05B2219/30—Nc systems
    • G05B2219/39—Robotics, robotics to robotics hand
    • G05B2219/39258—Three objective attitude control

Definitions

  • the present disclosure relates to an attitude determination and control system using a Quaternion Kalman Filter (QKF). It is more particularly to a satellite attitude control system compatible with a spherical motor as its attitude actuator. Moreover, the number of environmental sensors and the number of attitude actuators are expandable.
  • QKF Quaternion Kalman Filter
  • Satellites usually equip with payloads having pointing requirements, for instance, antennas, solar cells, or cameras.
  • these payloads are “fixed” on a satellite and cannot move or rotate by themselves. Therefore, they are only allowed to have linear or rotational movements together with the satellite.
  • the satellite has to perform an attitude compensation procedure which includes the following steps: (1) determining the attitude of the satellite and (2) adjusting the attitude of the satellite closer to the commanded attitude using actuators.
  • “Attitude” refers to the rotational relationship between an inertial reference frame and the reference frame adhered to the satellite. Moreover, they usually are expressed in the form of quaternions.
  • the inertial reference frame may be an Earth-Centered, Inertial (abbreviated as “ECI”) coordinate frame or an Earth-Centered, Earth-Fixed (abbreviated as “ECEF”) coordinate frame which is represented by the Cartesian coordinate system, the cylindrical coordinate system, the spherical coordinate system, or the like.
  • ECI Earth-Centered, Inertial
  • ECEF Earth-Fixed
  • Cartesian coordinate system can be defined by a given origin, an X-axis unit vector, and another Y-axis unit vector that is orthogonal to the X-axis unit vector.
  • the right-hand rule determines the Z-axis unit vector.
  • attitude measurement device (1) using an attitude measurement device or (2) using algorithms to perform calculations according to an environmental referencing object in the space like the sun or earth magnetic field.
  • the attitude can be directly obtained by using the attitude measurement device.
  • the attitude measurement device for example, may be a star tracker.
  • the attitude can also be obtained by using algorithms to perform calculations according to the environmental referencing objects measured by environmental sensors.
  • environmental sensors for example, may be sun sensors and magnetometers.
  • the environmental reference objects may be the sunlight or the earth's magnetic field.
  • the angular momentum stored in the angular momentum exchange device (such as a reaction wheel or a control moment gyro, namely the indirect attitude actuator) inside the satellite has to be known as an input of the attitude determination and control algorithms.
  • angular momentum stored in the satellite makes a difference in the power or the torque needed for a satellite to change its attitude.
  • an angular momentum exchange device of the satellite is functioning and storing a certain amount of angular momentum (hereinafter referred to as “1 st condition”) and 2) an angular momentum exchange device of the satellite is in the IDLE mode without storing any angular momentum (hereinafter referred to as “2 nd condition”).
  • the satellites require different torque to attend to the commanded attitude even the initial attitude is the same.
  • Such angular momentum stored has to be checked by rotational speed sensors in the angular momentum exchange device.
  • the conventional integrated satellite attitude determination and control system has several written attitude determination modes. However, each mode has an unchangeable number and type of 1) environmental sensors, 2) the attitude measurement unit, and 3) the inertia measurement unit. If the sensors designated by the mission requirement varied, the supplier must reprogram the attitude determination algorithm accordingly.
  • sensors may fail after a period of operation. If the system still solves for the attitude using the same algorithm, the system may face malfunctions.
  • an expandable algorithm with the capability to accommodate a different number of sensors is a solution for prolonging a satellite's lifespan.
  • the conventional attitude control system can't be expandable to have several environmental sensors.
  • they can't add or remove the inertia measurement unit (such as a gyroscope) or the attitude measurement unit (such as a star tracker) from the determination algorithm flexibly.
  • they also fail to add or remove different kinds of direct attitude actuators, indirect attitude actuators, or combinations thereof from the system.
  • the attitude determination algorithm has to be reprogrammed every time after the attitude measurement units, the direct attitude actuators and the indirect attitude actuators configuration is designated according to the client's requirements and budgets. Therefore, this traditional-style causes more delivery time of the system and the inconvenience of the client.
  • the algorithm can't be updated automatically, thereby causing a wrong determination of the satellite's attitude.
  • the object of the present disclosure is to provide an attitude control system based on a Quaternion Kalman Filter (QKF) and the method thereof.
  • the attitude control system mainly performs a calculation “according to the information of at least two ‘different’ environmental reference objects within the space measured by at least two environmental sensors” to achieve more accurate attitude determination and attitude control for the satellite.
  • the present disclosure provides an algorithm applied by the attitude control system, and the algorithm includes an attitude control and determination method.
  • the process of performing a calculation “according to the information of at least two different environmental reference objects within the space measured by at least two environmental sensors” includes at least one “propagate procedure,” and at least one “update procedure” to correct the propagation result.
  • the propagate procedure hereinafter is abbreviated as the “propagate stage,” on the other hand, the update procedure hereinafter is abbreviated as the “update stage.”
  • the attitude control method requires the attitude determination result for proceeding. Compared with the attitude determination method, the attitude control method at least further includes a “control procedure,” The control procedure hereinafter is abbreviated as the “control stage.”
  • the attitude determination algorithm includes one “propagate procedure,” and at least one “update procedure.” There is more than one kind of propagation method (namely, obtaining a first propagation result).
  • the present disclosure provides six of them. However, the present disclosure should not be limited to these six propagation methods. As long as the propagation methods are modified with the spirit of the present disclosure, they shall be encompassed by the scope of this present disclosure.
  • the attitude control system may have different elements or combinations of elements.
  • the attitude control system may further include an inertia measurement unit. So does the attitude information obtained by the inertia measurement unit.
  • the inertia measurement unit maybe, but is not limited to, a gyroscope, and the attitude measurement unit maybe, but is not limited to, a star tracker.
  • an attitude control system based on a QKF for controlling an attitude of a satellite.
  • the system includes a controller, an attitude actuator, a 1 st to N th environmental sensor, and a processor.
  • the controller receives an attitude command from a user to generate a torque-input command.
  • the attitude actuator is for outputting a torque according to a torque-control command to change the satellite's attitude.
  • a torque sensor can obtain a measured output torque from the attitude actuator.
  • the 1 st to the N th environmental sensor measures a 1 st to M th target to obtain a 1 st to N th measured value of environmental sensor (measured value).
  • N could be any positive integer
  • M is a positive integer less than or equal to N.
  • the processor has a QKF, and the processor connects to the controller, the attitude actuator, and the 1 st to the N th environmental sensor.
  • the processor obtains a first propagation result by calculating at least one selected from the group consisting of the torque-input command from the controller in the very last time step, and the measured output torque. For obtaining a second propagation result, the 1 st to the N th measured value is used for updating the first propagation result.
  • the attitude control system further includes at least one selected from the group consisting of an inertia measurement unit and an attitude measurement unit in some embodiments.
  • the inertia measurement unit measures the inertia of the satellite to obtain inertia information
  • the attitude measurement unit measures the satellite's attitude to obtain attitude information.
  • the processor is further connected to the inertia measurement unit, and the processor is for obtaining the first propagation result by performing a calculation according to at least one selected from the group consisting of “the torque-input command from the controller in the very last time step and the measured output torque,” and according to the “inertia information.”
  • the processor is further connected to the attitude measurement unit, and the processor is for obtaining the first propagation result by performing a calculation according to at least one selected from the group consisting of “the torque-input command from the controller in the very last time step and the measured output torque,” and according to the “attitude information.”
  • the attitude control system may include at least one selected from the group consisting of the “attitude measurement unit” and “the inertia measurement unit”. After the processor obtains the first propagation result, the processor further obtains the second propagation result by calculating the 1 st to the N th measured value and the first propagation result.
  • an attitude determination method based on a QKF for determining an attitude of a satellite includes a propagate stage and an update stage.
  • the propagate stage includes receiving an attitude command from a user and obtaining a first propagation result by using a processor having a QKF to perform a calculation according to at least one selected from the group consisting of the torque-input command from the controller in the very last time step and a measured output torque.
  • the update stage includes using a 1 st to N th environmental sensor to measure a 1 st to M th target to obtain a 1 st to N th measured value, wherein N could be any positive integer.
  • M is a positive integer less than or equal to N
  • the update stage includes obtaining a second propagation result by using the processor to perform a calculation according to the 1 st to the N th measured value and the first propagation result.
  • the propagate stage further includes at least one step selected from the group consisting of “using an inertia measurement unit to measure the inertia of the satellite”, and “using an attitude measurement unit to measure the satellite's attitude”.
  • the processor is for obtaining the first propagation result by performing a calculation according to at least one selected from the group consisting of “the “torque-input command from the controller in the very last time step and measured output torque,” and the inertia information.
  • the processor is for obtaining the first propagation result by performing a calculation according to at least one selected from the group consisting of “the torque-input command from the controller in the very last time step and the measured output torque,” and the attitude information.
  • the method includes a propagate stage, an update stage, and a control stage.
  • the propagate stage includes obtaining a first propagation result by using a processor having a QKF to perform a calculation according to at least one selected from the group consisting of “the torque-input command from the controller in the very last time step,” and “output torque measured by a torque sensor”.
  • the update stage includes using a 1 st to N th environmental sensor to measure a 1 st to M th target to obtain a 1 st to N th measured value, wherein N could be any positive integers.
  • This update stage includes obtaining a second propagation result using the processor to perform a calculation according to the 1 st to the N th measured value and updating the first propagation result.
  • the control stage calculates the torque-control command according to the “attitude command given by the user,” and the “second propagation result.” This torque-control command is then sent to triggering the attitude actuator for changing the attitude of a satellite.
  • the propagate stage further includes at least one step selected from the group consisting of “using an inertia measurement unit to measure an inertia of the satellite to obtain an inertia information” and “using an attitude measurement unit to measure the attitude of the satellite to obtain an attitude information”.
  • the propagate stage further includes a step of “using the inertia measurement unit to measure the inertia of the satellite to obtain the inertia information”
  • the processor is for obtaining the first propagation result by performing a calculation according to at least one selected from the group consisting of “the torque command from the processor in the very last time step and the measured output torque,” and the “inertia information.”
  • the processor obtains the first propagation result according to at least one selected from the group consisting of “the input-torque command from the controller in the very last time step and the measured output torque,” and the “attitude information.”
  • the propagate stage includes “using an inertia measurement unit to measure an inertia of the satellite to obtain an inertia information” and “using an attitude measurement unit to measure the attitude of the satellite to obtain an attitude information”.
  • the processor is for obtaining the first propagation result by performing a calculation according to the “input-torque command from the controller in the very last time step,” the “measured output torque,” the “inertia information,” and the “attitude information”.
  • the attitude actuator is a direct attitude actuator.
  • the attitude actuator further includes at least one direct attitude actuator and at least one indirect attitude actuator.
  • the indirect attitude actuator could be a spherical motor, and the controller of the control system has at least three control modes: a de-tumbling mode, a de-saturation mode, and a fine pointing mode.
  • FIG. 1 illustrates a schematic view of a satellite attitude control system E 11 according to an embodiment of the present disclosure
  • FIG. 2 illustrates a schematic view of a satellite attitude control system E 11 according to another embodiment of the present disclosure
  • FIG. 3 illustrates a schematic view of a satellite attitude control system E 12 according to another embodiment of the present disclosure
  • FIG. 4 illustrates a schematic view of a satellite attitude control system E 13 according to another embodiment of the present disclosure
  • FIG. 5 illustrates a flowchart of satellite attitude determination methods E 21 -E 26 and a satellite attitude control method E 31 according to one embodiment of the present disclosure
  • FIG. 6 illustrates a flowchart of a propagate stage in the satellite attitude determination methods E 21 -E 22 according to another embodiment of the present disclosure
  • FIG. 7 illustrates a flowchart of the propagate stage in the satellite attitude determination method E 21 according to another embodiment of the present disclosure
  • FIG. 8 illustrates a flowchart of the propagate stage in the satellite attitude determination method E 22 according to another embodiment of the present disclosure
  • FIG. 9 illustrates a flowchart of the propagate stage in the satellite attitude determination methods E 23 -E 27 according to one embodiment of the present disclosure
  • FIG. 10 illustrates a flowchart of the propagate stage in the satellite attitude determination method E 23 according to one embodiment of the present disclosure
  • FIG. 11 illustrates a flowchart of the propagate stage in the satellite attitude determination method E 24 according to one embodiment of the present disclosure
  • FIG. 12 illustrates a flowchart of the propagate stage in the satellite attitude determination method E 25 according to one embodiment of the present disclosure
  • FIG. 13 illustrates a flowchart of the propagate stage in the satellite attitude determination method E 26 according to one embodiment of the present disclosure
  • FIG. 14 illustrates a flowchart of the propagate stage in the satellite attitude determination method E 27 according to one embodiment of the present disclosure
  • FIG. 15 illustrates a flowchart of an update stage in the satellite attitude determination methods E 21 -E 27 according to one embodiment of the present disclosure.
  • FIG. 16 illustrates a flowchart of a control stage in the satellite attitude control method E 31 according to one embodiment of the present disclosure.
  • the directional indications (such as front, rear, left, right, above, under, clockwise, counterclockwise, etc.) in the embodiments are only used to describe the relative positions or movements, etc. of multiple components under specific circumstances (such as those shown in the drawings). When the circumstances are changed, the directional indications may be changed accordingly.
  • the terms “one embodiment,” “one practical embodiment,” “a preferred embodiment,” “the other embodiment,” “another embodiment,” or any descriptions which are related to the “embodiment” or the “implementation” are provided for describing the feature, the structure, or the property that may be included in the described embodiment; it should be noted that other embodiments do not necessarily include the feature, the structure, or the property. Moreover, using the terms “in this embodiment,” “in some embodiments,” or the like does not limit the described embodiments to be the same embodiment. It may refer to different embodiments.
  • the conventional QKF cannot adjust its number of environmental sensors and actuators without reprogramming its software or firmware.
  • a technical feature is applying the QKF on the attitude control system. It could be applied on a satellite, an underwater vehicle, a drone, etc. Detail operations of the QKF may be referred to D. Choukroun, L. Y. Bar-Itzhack, and Y. Oshman. (2005) Novel Quaternion Kalman Filter. Published on IEEE Transactions on Aerospace and Electronic Systems Vol. 42, No. 1 (abbreviated as “Reference 1”).
  • the algorithm built in the satellite attitude control system is improved, thereby completing the system's architecture being expandable to have several environmental sensors.
  • the Kalman filter is a high-efficiency recursive filter (autoregressive filter).
  • the Kalman filter uses a series of measurements containing random errors or system errors, produces estimates of the state of a dynamic system.
  • the algorithm built in the satellite attitude control system includes three parts: initialize (hereinafter, “Part A”), propagate (hereinafter, “Part B”), and update (hereinafter, “Part C”).
  • the following parameters have to be defined in advance.
  • Part A The calculation processes of Part A, Part B, and Part C are described in order.
  • the relationship between the attitude quaternion (namely, q k/k ) of the satellite and the time k and the relationship among the state space ( ⁇ circumflex over (x) ⁇ k/k T ) of the satellite, the attitude quaternion ( ⁇ circumflex over (q) ⁇ k/k T ) of the satellite, and the drift estimations ( ⁇ circumflex over ( ⁇ ) ⁇ k/k T ) at time k can be presented by following Eq. A-1 and Eq. A-2:
  • Part A Initialize
  • Part B Propagate
  • Part C Update
  • the algorithm provided by the present disclosure can support infinite environmental sensors, namely, infinite times of updates (the Part C cycles).
  • “one” propagate and update cycle may include one propagate and one update, or one propagate and two updates, or one propagate and three updates, or the like.
  • the satellite attitude control system has N environmental sensors which are installed and participate in the tasks of the satellite attitude control system (where both N and M are integers greater than or equal to 2) actually, then the “one” propagate and update cycle includes one propagate and N updates.
  • the dynamic equations and the predecessor works of the satellite system of the embodiments of the present disclosure are described in advance. That is, obtaining the satellite angular acceleration (a k ) of Eq. B-1 and Eq. B-2; the satellite angular acceleration may be presented by the first derivative of the satellite angular velocity ( ) After the first derivative of the satellite angular velocity ( ) is obtained, the propagate of the Kalman filter is performed.
  • a governing equation is applied to calculate and express the relationship between the angular momentum and the torque.
  • the governing equation considers the direct actuator, the indirect actuator, and the disturbance torque.
  • Eq. B0-1′, Eq. B0-1′′, and Eq. B0-1 are substantially the same.
  • the shifts of the items are the equations is for the sake of convenience in describing and expressing the conservation law for angular momentum in the embodiments of the present disclosure.
  • Eq. B0-4 and Eq. B0-5 are results develop from some items in Eq. B0-1.
  • the development method may be referred to GEORGE W. HOUSNER, DONALD E. HUDSON. (1991) Applied Mechanics Dynamics (abbreviated as Reference 2), p 202-p 203, Section 7.8 “The General Equations of Motion for a Rigid Body.”
  • the model describes the system dynamics of any indirect actuator, including a spherical motor (e.g., a motor disclosed in Taiwan Invention Patent (Patent Number I719585) “Motor and driving method thereof”), a single-gimbal control moment gyro, a dual-gimbal control moment gyro, a reaction wheel, etc.
  • the product of inertia (POI) of the rotor and the satellite itself is fully considered in the system model.
  • H satellite ⁇ I satellite ⁇ ⁇ satellite ⁇ + ⁇ satellite ⁇ H satellite ( Eq . B0 - 3 )
  • H indirect - actuator ⁇ ( a ) I rotor ⁇ ( a ) ⁇ ⁇ rotor ⁇ ( a ) ( Eq . B0 - 4 )
  • H indirect - actuator ⁇ ( a ) ⁇ I rotor ⁇ ( a ) ⁇ ⁇ rotor ⁇ ( a ) ⁇ + ⁇ rotor ⁇ ( a ) ⁇ H rotor ⁇ ( a ) ( Eq . B0 - 5 )
  • the satellite attitude control system can, by measuring, inferring, and calculating, directly or indirectly obtain the rotational speed of the rotor ⁇ rotor(a) , the angular acceleration of the rotor ⁇ , the inertia tensor of the rotor I rotor(a) , the angular velocity of the satellite ⁇ satellite , and the inertia tensor of the satellite I satellite . Therefore, from the foregoing equations, to can be derived as following Eq. B0-6.
  • the satellite attitude control system of the embodiment of the present disclosure is expandable to have any number of direct attitude actuators and indirect actuators freely.
  • the indirect actuators may have several implementations, these indirect actuators are all based on the conservation law of angular momentum, as indicated by Eq. B0-1.
  • the indirect actuators in different implementations are all suitable for being applied in the satellite attitude control system of the present disclosure.
  • the corrected delta angle estimation ( ) is introduced into Eq. B-4 to Eq. B-6 for the calculation to obtain a propagated state space at time k+1 ( ⁇ circumflex over (x) ⁇ k+1/k ), the ⁇ circumflex over (x) ⁇ k+1/k is also “the first attitude propagation result” in any of the following embodiments.
  • the “ ⁇ ” represents the ⁇ function. Detailed contents of the ⁇ function can be referred to page 176 of Reference 1.
  • the aforementioned corrected delta angle estimation ( ) is introduced into following Eq. B-7 to Eq. B-13 for the calculation to obtain corrected error covariance matrix ( ).
  • the satellite attitude control system has N environmental sensors which are installed and participate in the update stage (where both N and M are integers greater than or equal to 2).
  • N here stands for the designated number of the environmental sensor, which is a number ranging from 0 to N (where both N and M are integers greater than or equal to 2). It is understood that “N” not only represents the number of environmental sensors on the satellite but also represents the times of “update” in the algorithm.
  • Part C the characteristics of the quaternion are utilized to obtain an H function, and when the H function multiplies a quaternion which can correctly map b k+1 i to r k+1 i , the product of the H function and the quaternion equal 0.
  • the Kalman Gain is calculated by introducing the H function and values at k+1 obtained in Part B (Propagate), which include the state space ⁇ circumflex over (x) ⁇ k+1/k , the error covariance matrix P k+1/k , the environment sensing vector pair [b k+1 i , r k+1 i ] (where b k+1 i is the measured value of the environmental sensor) into following Eq. C-05 to Eq. C-09.
  • Eq. C-10 and Eq. C-11 are applied to “update” the system state and the error covariance matrix.
  • k A particular time step that the controller received attitude command from the user and commanded the attitude actuator.
  • k + 1 The time step that follows right after the time step k ⁇ k “Measured output torque” or “torque-control command” ⁇ circumflex over (x) ⁇ k + 1/k First propagation result b k + 1 i
  • the measured value of the environmental sensor N The satellite attitude control system has N environmental sensors which are installed and participated in the tasks of the satellite attitude control system i An integer ranging from 0 to N representing the designated number of certain environmental attitude sensors Second propagation result
  • FIG. 1 and FIG. 2 illustrate schematic views of a satellite attitude control system E 1 according to a first embodiment of the present disclosure.
  • the satellite attitude control system is operated based on a Quaternion Kalman Filter 41 (QKF 41).
  • QKF 41 Quaternion Kalman Filter 41
  • the satellite attitude control system based on the Quaternion Kalman Filter is provided for controlling an attitude of a satellite.
  • the system includes a controller 1 , an attitude actuator 2 , a 1 st to N th environmental sensor, and a processor 4 .
  • Controller 1 is for receiving an attitude command of the satellite requested by a user to generate a torque-input command.
  • the attitude actuator 2 is for outputting a torque according to a torque-control command to change the satellite's attitude.
  • the attitude actuator 2 can be measured by a torque sensor to obtain a measured output torque value.
  • the 1 st to the N th environmental sensor measures a 1 st to M th target to obtain a 1 st to N th measured value. Both N and M are integers greater than or equal to 2, and M is a positive integer less than or equal to N.
  • Processor 4 has a QKF 41, and processor 4 is connected to controller 1 , the attitude actuator 2 , and the 1 st to the N th environmental sensor. Processor 4 performs a calculation according to at least one selected from the group consisting of “the torque-input command from the controller in the very last time step and the measured output torque” to obtain a first propagation result.
  • processor 4 calculates the second propagation result according to the 1 st to the N th measured value via updating the first propagation result.
  • processor 4 is further for performing a calculation for correction purposes according to the “attitude command,” and the “second propagation result” to obtain a torque-control command, and outputting torque to the satellite according to the torque-control command to change the attitude of the satellite by using the attitude actuator 2 .
  • the satellite could be, but not limited to, a satellite or an underwater vehicle.
  • the satellite attitude control system is applied to a satellite as an illustrative example in the following embodiments.
  • the satellite attitude control system is applied to execute (1) determine the attitude of the satellite at time t, and (2) controlling and adjusting the attitude of the satellite closer to the commanded attitude at time t+1.
  • the attitude command can be represented by quaternions or the like.
  • the 1 st to the N th environmental sensor is the first environmental sensor 31 to the N th environmental sensor 3 N.
  • N is equal to 2.
  • the satellite attitude control system E 11 includes the first environmental sensor 31 and the second environmental sensor 32 .
  • N is greater than 2.
  • the satellite attitude control system E 11 includes the first environmental sensor 31 , the second environmental sensor 32 , and the N th environmental sensor 3 N.
  • FIG. 1 to FIG. 4 illustrates that the satellite attitude control systems E 11 , E 12 , E 13 , according to embodiments of the present disclosure, are expandable to have N environmental sensors.
  • processor 4 calculates at least one input from environmental sensors that obtains measured value in the satellite body reference frame.
  • the ideal measured value for the same environmental referencing object in the inertia frame can further be calculated via ephemeris or any celestial model. By comparing these two pieces of information, a correction for the first propagation result could be made.
  • an algorithm capable of performing calculation or data processing to two measured values e.g., the position of the sun and the position of the earth
  • two measured values e.g., the position of the sun and the position of the earth
  • processor 4 of the satellite attitude control system is connected to two environmental sensors
  • processor 4 is connected to two different types of environmental sensors (e.g., the sun sensor and the magnetometer) to measure two “different” targets (for example, but not limited to the sun and the earth magnet field).
  • the calculation for the number of the environmental sensors is neither based on the number of the devices nor based on the maximum number of the environmental sensors that can be disassembled or divided without damage, and it is not calculated based on the types of the environmental sensor. Instead, the number of the environmental sensors is calculated based on the number of the smallest units that can individually and effectively measure the target.
  • the satellite attitude control system may use five sun sensors to measure the position of the sun (the first target) and one earth magnetic sensor to measure the position of the earth (the second target) at the same time.
  • the number of the environmental sensors is calculated based on the number of different effective measurements of one environmental sensor in the propagate stage S 1 and the update stage S 2 .
  • one environmental sensor can sequentially or simultaneously perform measurements to the sun and the magnetic field of the earth
  • the environmental sensor in the satellite attitude determination method performs effective measurements for the two different targets (the position of the sun and the magnetic field of the earth) sequentially or simultaneously, it is considered that two environmental sensors are applied in this scenario.
  • the satellite attitude control system is expandable to have two or more environmental sensors.
  • the satellite attitude control system can simultaneously connect to N environmental sensors, respectively named the 1 st to N th environmental sensor, which can sequentially or simultaneously measure M “different” targets, respectively named the 1 st to the N th measured value.
  • N environmental sensors respectively named the 1 st to N th environmental sensor
  • M “different” targets respectively named the 1 st to the N th measured value.
  • the satellite attitude control system uses six environmental sensors; five of the six environmental sensors are sun sensors, the target corresponding to these five environmental sensors should be the sun, the rest one of the six environmental sensors is a magnetometer, and the target corresponding to these environmental sensors should be the earth. Therefore, N equals 6, and m equals 2.
  • the satellite attitude control system has at least two functions capable of obtaining a second propagation result, but not limited to.
  • the satellite attitude control system may have the following options: (1) obtaining a first propagation result by performing a calculation according to at least one selected from the group consisting of the torque-input command in the very last time step, and obtaining a second propagation result by performing a calculation according to the 1 st to the N th measured value and updating the first propagation result; or (2) obtaining a first propagation result by performing a calculation according to the measured value of the torque output, and obtaining a second propagation result by performing a calculation according to the 1 st to the N th measured value and updating the first propagation result.
  • the attitude actuator 2 is a direct attitude actuator 2 , for example, a magnetorquer or a thruster.
  • the attitude actuator 2 further includes at least one direct attitude actuator and at least one indirect attitude actuator.
  • the indirect attitude actuator may include but not be limited to a spherical motor (e.g., a motor disclosed in Taiwan Invention Patent (Patent Number I719585) “Motor and driving method thereof”), a single-gimbal control moment gyro, a dual-gimbal control moment gyro, a reaction wheel, etc.
  • the spherical motor needs to have the “angular acceleration (unit: rad/s 2 , which is represented as “ ⁇ ”).”
  • the spherical motor has limits on the rotational speed, and the spherical motor cannot be accelerated anymore when it hits its maximum rotational speed.
  • the spherical motor automatically enters into the De-saturation mode, in which a direct attitude actuator is applied to generate a reverse torque whose direction is opposite to the direction of the torque generated by the indirect attitude actuator while deacceleration.
  • This reverse torque is applied to compensate with the torque produce by the spherical motor while deacceleration, thereby achieving the purpose of the De-saturation for the satellite.
  • the attitude actuator 2 in the satellite attitude control system can be solely the direct attitude actuator 2 but cannot be the indirect attitude actuator 2 .
  • FIG. 3 illustrates a schematic view of a satellite attitude control system E 12 according to another embodiment of the present disclosure.
  • the satellite attitude control system E 12 further includes an inertia measurement unit 5 .
  • the inertia measurement unit 5 is for measuring the inertia of the satellite to obtain inertia information.
  • the processor 4 is further connected to the inertia measurement unit 5 .
  • Processor 4 obtains the first propagation result by performing a calculation according to “at least one selected from the group consisting of the torque-input command from the controller in the very last time step” and the measured output torque,” and the “inertia information.”
  • processor 4 of the satellite attitude control system E 12 obtains the first propagation result by performing a calculation according to the “the torque-input command from the controller in the very last time step,” and the “inertia information.”
  • processor 4 of the satellite attitude control system E 12 obtains the first propagation result by calculating the “measured output torque,” and the “inertia information.”
  • FIG. 4 illustrates a schematic view of a satellite attitude control system E 13 according to another embodiment of the present disclosure.
  • the satellite attitude control system E 13 further includes an attitude measurement unit 6 .
  • the attitude measurement unit 6 is for measuring the attitude of the satellite to obtain attitude information.
  • the processor 4 is further connected to the attitude measurement unit 6 .
  • Processor 4 obtains the first propagation result by performing a calculation according to “at least one selected from the group consisting of the torque-input command from the controller in the very last time step and the measured output torque,” and the “attitude information.”
  • processor 4 of the satellite attitude control system E 13 obtains the first propagation result by performing a calculation according to the “the torque-input command from the controller in the very last time step,” and the “attitude information.”
  • processor 4 of the satellite attitude control system E 13 obtains the first propagation result by performing a calculation according to the “measured output torque,” and the “attitude information.”
  • the fourth to the tenth practical embodiments (satellite attitude determination methods E 21 -E 26 ) as well as the eleventh practical embodiment (satellite attitude control method E 31 ) are provided to elaborate the determination and control method for satellite attitude implemented by the foregoing satellite attitude control system of the embodiment of the present disclosure.
  • the major differences between the practical embodiments are described as below.
  • FIG. 5 illustrates a flowchart of satellite attitude determination methods E 21 -E 26 according to one embodiment of the present disclosure and a satellite attitude control method E 31 according to one embodiment of the present disclosure. It should be noted that as compared with the satellite attitude determination methods E 21 -E 26 according to the present disclosure, the satellite attitude control method E 31 according to the present disclosure merely further includes the “control stage S 3 ”.
  • the satellite attitude determination methods according to the present disclosure may be mainly divided into two types (six kinds in sum):
  • FIG. 6 illustrates a first type of the satellite attitude determination method according to the present disclosure, which includes the satellite attitude determination methods E 21 and E 22 , and the satellite attitude determination methods E 21 and E 22 are respectively described by the fourth practical embodiment and the fifth practical embodiment;
  • FIG. 9 illustrates a second type of the satellite attitude determination method according to the present disclosure, which includes the satellite attitude determination methods E 23 to E 26
  • the sixth practical embodiment describes the satellite attitude determination methods E 23 to E 26 to the ninth practical embodiment.
  • the difference between the second type and the first type is that the second type of the satellite attitude determination methods E 23 to E 26 , as compared with the first type of the satellite attitude determination methods E 21 and E 22 , further includes “measuring the inertia of the satellite to obtain an inertia information by using an inertia measurement unit 5 ” (step S 12 a ) or “measuring the attitude of the satellite to obtain an attitude information by using an attitude measurement unit 6 ” (step S 12 b ).
  • the second type of satellite attitude determination method includes the satellite attitude determination methods E 23 to E 26 .
  • the difference between these satellite attitude determination methods E 23 to E 26 is that, in the propagate stage, according to which “parameters,” the processor 4 having the QKF 41 performs the calculation to obtain the first propagation result (step S 15 ).
  • the “parameters” may be:
  • FIG. 7 illustrates a flowchart of a propagates stage S 1 in the satellite attitude determination method E 21 according to another embodiment of the present disclosure.
  • Another embodiment of the present disclosure provides a satellite attitude determination method based on the QKF for predicting the satellite's attitude.
  • the method includes a propagate stage S 1 and an update stage S 2 .
  • the propagate stage S 1 includes:
  • FIG. 8 illustrates a flowchart of a propagates stage S 1 in the satellite attitude determination method E 22 according to another embodiment of the present disclosure.
  • Another embodiment of the present disclosure provides a satellite attitude determination method based on the QFK for predicting the satellite's attitude.
  • the method includes a propagate stage S 1 and an update stage S 2 .
  • the propagate stage S 1 includes:
  • FIG. 9 illustrates a flowchart of a propagates stage S 1 in the satellite attitude determination methods E 23 , E 24 , E 25 , E 26 according to one embodiment of the present disclosure.
  • the differences between these satellite attitude determination methods E 23 to E 26 is that, in the propagate stage, according to which “parameters,” the processor 4 having the QKF performs the calculation to obtain the first propagation result (step S 15 ).
  • the “parameters” may be:
  • FIG. 10 illustrates a flowchart of a propagates stage S 1 in the satellite attitude determination method E 23 according to one embodiment of the present disclosure.
  • the propagate stage S 1 in the satellite attitude determination method E 23 further includes:
  • FIG. 11 illustrates a flowchart of a propagates stage S 1 in the satellite attitude determination method E 24 according to one embodiment of the present disclosure.
  • the propagate stage S 1 in the satellite attitude determination method E 24 further includes:
  • FIG. 12 illustrates a flowchart of a propagates stage S 1 in the satellite attitude determination method E 25 according to one embodiment of the present disclosure.
  • the propagate stage S 1 in the satellite attitude determination method E 25 further includes:
  • FIG. 13 illustrates a flowchart of a propagates stage S 1 in the satellite attitude determination method E 26 according to one embodiment of the present disclosure.
  • the propagate stage S 1 in the satellite attitude determination method E 26 further includes:
  • FIG. 14 illustrates a flowchart of a propagates stage S 1 in the satellite attitude determination method E 27 according to one embodiment of the present disclosure.
  • the propagate stage S 1 in the satellite attitude determination method E 27 further includes:
  • FIG. 15 illustrates a flowchart of an update stage S 2 in the satellite attitude determination methods E 21 -E 26 according to one embodiment of the present disclosure. It should be noted that, though the steps of the propagate stages S 1 in these satellite attitude determination method E 21 -E 26 are different, these satellite attitude determination method E 21 -E 26 have the same update stage S 2 after obtaining the first propagation result, and the update stage includes:
  • FIG. 16 illustrates a flowchart of a control stage S 3 in a satellite attitude control method E 31 according to one embodiment of the present disclosure.
  • the satellite attitude control method E 31 according to the present disclosure merely further includes the “control stage S 3 ”.
  • the satellite attitude control method includes any of the satellite attitude determination methods E 21 -E 26
  • the control stage S 3 includes:
  • the satellite attitude control method E 31 described in the foregoing paragraphs may be generally called “the satellite fine pointing mode.”
  • the satellite attitude control system compares the “attitude command” from the user with the second attitude propagation result to obtain the torque control command for the attitude actuator. Consequently, the satellite's attitude may reach its “attitude command” one step closer.
  • the satellite attitude control method E 31 includes at least three control modes. In addition to the foregoing satellite fine pointing mode, the satellite attitude control method E 31 further includes a satellite de-tumbling mode (De-tumbling) and a satellite de-saturation mode (De-Saturation).
  • a satellite de-tumbling mode (De-tumbling)
  • a satellite de-saturation mode (De-Saturation).
  • the attitude control method E 31 further includes the satellite de-saturation mode (De-saturation).
  • the De-tumbling mode is an emergency mode triggered when (1) satellites just get ejected into space from the rocket or (2) the satellite is suffering from significant disturbance torque and entering a tumbling state.
  • the satellite detects this tumbling state happens herein refers to the absolute body angular velocity of the satellite is faster than a preset value, the de-tumbling mode is triggered.
  • the direct attitude actuator 2 is used in the de-tumbling mode. It outputs torque to the satellite according to the torque-control command to reduce the tumbling speed of the satellite.
  • a satellite attitude control system based on the QKF and control method thereof are provided.
  • the satellite attitude control system can achieve the effects of “expandable environmental sensors,” “performs a calculation according to the information of ‘different’ environmental reference objects within the space measured by environmental sensors,” and “performs at least one ‘update procedure’ to correct the propagation result” to the first propagation result.
  • an “expandable attitude determination system” can be achieved.
  • the present disclosure provides a satellite attitude control method in which various types and numbers of indirect attitude actuators (e.g., a spherical motor) and direct attitude actuators can be applied.
  • indirect attitude actuators e.g., a spherical motor
  • direct attitude actuators can be applied.

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Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6285927B1 (en) * 1999-05-26 2001-09-04 Hughes Electronics Corporation Spacecraft attitude determination system and method
US7246775B1 (en) * 2004-08-02 2007-07-24 Lockheed Martin Corporation System and method of substantially autonomous geosynchronous time-optimal orbit transfer
US9038958B1 (en) * 2012-05-29 2015-05-26 United States Of America As Represented By The Secretary Of The Navy Method and apparatus for contingency guidance of a CMG-actuated spacecraft
US20200255165A1 (en) * 2019-02-12 2020-08-13 Canadian Space Agency Spacecraft control using residual dipole

Family Cites Families (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6477465B1 (en) * 1999-11-29 2002-11-05 American Gnc Corporation Vehicle self-carried positioning method and system thereof
US7193559B2 (en) * 2003-01-21 2007-03-20 Novatel, Inc. Inertial GPS navigation system with modified kalman filter
US20110238308A1 (en) * 2010-03-26 2011-09-29 Isaac Thomas Miller Pedal navigation using leo signals and body-mounted sensors
US9194756B2 (en) * 2010-07-14 2015-11-24 University Of Florida Research Foundation, Inc. System and method for assessing the performance of an attitude control system for small satellites
TW201311214A (zh) * 2011-09-14 2013-03-16 Ming-Da Lin 視覺化ｉｍｕ步態偵測裝置及其分析方法
US9475592B2 (en) * 2013-01-31 2016-10-25 Northrop Grumman Systems Corporation Reaction sphere for stabilization and control in three axes
AU2015222926A1 (en) * 2014-02-26 2016-10-13 Clark Emerson Cohen An improved performance and cost Global Navigation Satellite System architecture
TW201709025A (zh) * 2015-08-26 2017-03-01 巨擘科技股份有限公司 整合位置、姿態與無線傳輸之裝置
CN113804191B (zh) * 2016-11-17 2024-03-19 格兰菲智能科技有限公司 移动设备及求取移动设备姿态的方法
US10371530B2 (en) * 2017-01-04 2019-08-06 Qualcomm Incorporated Systems and methods for using a global positioning system velocity in visual-inertial odometry
WO2020006667A1 (en) * 2018-07-02 2020-01-09 Beijing DIDI Infinity Technology and Development Co., Ltd Vehicle navigation system using pose estimation based on point cloud
TWI680382B (zh) * 2018-10-19 2019-12-21 宏達國際電子股份有限公司 電子裝置及其姿態校正方法

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6285927B1 (en) * 1999-05-26 2001-09-04 Hughes Electronics Corporation Spacecraft attitude determination system and method
US7246775B1 (en) * 2004-08-02 2007-07-24 Lockheed Martin Corporation System and method of substantially autonomous geosynchronous time-optimal orbit transfer
US9038958B1 (en) * 2012-05-29 2015-05-26 United States Of America As Represented By The Secretary Of The Navy Method and apparatus for contingency guidance of a CMG-actuated spacecraft
US20200255165A1 (en) * 2019-02-12 2020-08-13 Canadian Space Agency Spacecraft control using residual dipole

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