WO2021068554A1 - 一种低轨道地磁蓄能在轨投送的航天器章动抑制方法 - Google Patents

一种低轨道地磁蓄能在轨投送的航天器章动抑制方法 Download PDF

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WO2021068554A1
WO2021068554A1 PCT/CN2020/097729 CN2020097729W WO2021068554A1 WO 2021068554 A1 WO2021068554 A1 WO 2021068554A1 CN 2020097729 W CN2020097729 W CN 2020097729W WO 2021068554 A1 WO2021068554 A1 WO 2021068554A1
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Prior art keywords
delivery
spacecraft
link
space target
orbit
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PCT/CN2020/097729
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English (en)
French (fr)
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李文皓
张珩
冯冠华
张琛
杨磊
吕林立
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中国科学院力学研究所
中国科学院力学研究所广东空天科技研究院
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Priority to US17/638,870 priority Critical patent/US11530054B2/en
Publication of WO2021068554A1 publication Critical patent/WO2021068554A1/zh

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    • 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
    • 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
    • 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
    • B64G1/242—Orbits and trajectories
    • 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
    • B64G1/38—Guiding or controlling apparatus, e.g. for attitude control damping of oscillations, e.g. nutation dampers
    • 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
    • B64G1/64—Systems for coupling or separating cosmonautic vehicles or parts thereof, e.g. docking arrangements
    • 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
    • B64G1/64—Systems for coupling or separating cosmonautic vehicles or parts thereof, e.g. docking arrangements
    • B64G1/646—Docking or rendezvous 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
    • B64G1/66—Arrangements or adaptations of apparatus or instruments, not otherwise provided for

Definitions

  • the invention relates to the technical field of low-orbit geomagnetic energy storage on-orbit delivery.
  • the purpose of the present invention is to provide a spacecraft nutation suppression method for low-orbit geomagnetic energy storage on-orbit delivery. Before and after several state changes in the preparation process of geomagnetic energy storage rotation, energy dissipation and unloading, re-projection, Effectively restrain the free nutation ability of the spacecraft system in orbit.
  • a spacecraft nutation suppression method for on-orbit delivery of low-orbit geomagnetic energy storage includes the following control steps:
  • the spacecraft system is divided into the spacecraft body one and the two spacecraft body which are fixedly connected by the main connecting shaft.
  • the main connecting shaft is rotated with a projection link perpendicular to it, and the projection link slides along the length direction.
  • Two mass blocks are connected, and the center of mass of the spacecraft system is adjusted through the main connecting shaft;
  • the space target or de-orbit debris remains in the corresponding position of the delivery link, and the center of mass of the delivery link of the space target or de-orbit debris to be projected and
  • the main axis of inertia is measured, calibrated and adjusted respectively.
  • the adjusted main connecting axis around the projecting link passes through the center of mass of the spacecraft system after grabbing the space target or de-orbiting debris, and is in line with the rotation direction of the projecting link.
  • the main axes of inertia overlap, and the projecting plane of the projecting link that rotates vertically around the main connecting axis passes through the center of mass of the spacecraft system after grabbing the space target or de-orbiting debris;
  • the unloading process is the inverse process of energy storage and delivery, and the energy dissipating unloads the moment of inertia of the delivery link that rotates vertically around the main connecting axis until it stops rotating;
  • the spacecraft system is ready to grab the next space target or off-orbit debris and enter the next delivery cycle.
  • the spacecraft system adopts a split structure, and the main connecting shaft is connected with a linear telescopic mechanism for measuring whether the spacecraft system's center of mass passes through the main connecting shaft.
  • the projection connecting rod is slidably connected in the length direction.
  • the end of the delivery link is equipped with a cage for holding space targets or off-orbit debris, and the sequence of attitude rotation measurement and center of mass adjustment is reasonably allocated, so that the geomagnetic energy storage rotates and delivers-dissipates Can be unloaded-before and after the sudden change of several states in the process of re-projection preparation, the center of mass of the spacecraft system is always at the intersection of the main connecting axis and the sudden moment of inertia plane, thereby effectively preventing the spacecraft system from failing due to the rotating inertia main axis
  • the center of mass causes nutation problems.
  • step S1 is further configured as follows: the specific adjustment steps of step S1 are as follows:
  • the linear telescopic device connected to the main connecting shaft performs telescopic work, and the relative position between the spacecraft body 1 and the spacecraft body 2 connected at the two ends of the main connecting shaft is extended or contracted, and the expansion and contraction of the linear telescopic device is measured.
  • the attitude rotation of the spacecraft system changes.
  • the center of mass of the spacecraft system does not pass through the main connecting axis;
  • step (3) Adjust the mass distribution in the spacecraft body one and the spacecraft body two, and repeat step (2) until the spacecraft system no longer changes in attitude rotation during the expansion and contraction of the linear telescopic device, and the adjustment of the spacecraft is completed.
  • the center of mass of the system passes through the main connecting shaft;
  • the linear telescopic device performs telescopic operations until step (4) is repeated in a certain telescopic state of the linear telescopic device, and the spacecraft system no longer changes in attitude rotation, and the center of mass of the spacecraft system is located on the main connecting axis at the same time.
  • step (4) is repeated in a certain telescopic state of the linear telescopic device, and the spacecraft system no longer changes in attitude rotation, and the center of mass of the spacecraft system is located on the main connecting axis at the same time.
  • the delivery link rotates vertically around the main connection shaft, slide the two masses on the delivery link back to the main connection shaft; that is, under no-load conditions, calibrate the expansion and contraction of the linear telescopic device
  • the status is the corresponding no-load 0 position of the delivery link;
  • step S2 is further configured as follows: the specific adjustment steps of step S2 are as follows:
  • the spacecraft system grabs the space target or off-orbit debris to be projected, keeps the space target or off-orbit debris in the corresponding position of the projecting link, and slides two masses along the length of the projecting link. Until the linear telescopic device is telescopic, the measuring spacecraft system no longer changes in attitude rotation, that is, the balance position of the two masses on the delivery link of the space target to be delivered or the off-orbit debris.
  • the mass center of the delivery link of the space target or off-orbit debris to be delivered passes through the main connecting shaft;
  • the balance position on the delivery link of the target or de-orbit debris is the optimal position determined by the two masses of the space target or de-orbit debris to be delivered, and it is calibrated after the spacecraft system grabs the space target or de-orbit debris.
  • the center of mass and the inertia principal axis of the projecting link of the space target or off-orbit debris to be projected are measured, calibrated and adjusted respectively, so that the main connecting shaft around the projecting link can grasp the space target or The center of mass of the spacecraft system after de-orbiting debris, and overlaps with the inertia principal axis of the rotation direction of the projection link, and the projection plane where the projection link rotates vertically around the main link axis is used to grasp the space target or de-orbit The center of mass of the spacecraft system after the debris.
  • the present invention is further configured that: the attitude rotation change of the spacecraft system is a pitch, yaw, or roll angle change.
  • the present invention is further configured as follows: when the main connecting shaft is rotated and provided with a projecting link perpendicular to it, the projecting link adopts a geomagnetic energy storage method to accelerate the rotation of the projecting link in step S3.
  • the orthogonal strong magnetic moment generating device is unidirectionally affected by the external torque of the earth’s magnetic field and the transmission support that reacts to the torque transmission mechanism.
  • the internal moments of the rotating parts are balanced, and the spacecraft system will not accelerate the rotation of the attitude during the energy storage acceleration process.
  • the present invention is further configured as follows: when the main connecting shaft is rotatably provided with two projecting linkages perpendicular to it, in the step S3, the two projecting linkages are driven by a counter-rotating transmission mechanism to perform energy storage acceleration reaction. To rotate.
  • the acceleration mode of geomagnetic energy storage is not adopted, and the fixed support of the counter-rotating transmission mechanism is internally reacted by the positive rotating member.
  • the torque is balanced with the internal torque reacted by the counter-rotating member, and the spacecraft system will not accelerate the rotation of the attitude during the energy storage acceleration process.
  • step S4 the specific steps of adjusting the delivery link of the space target or off-orbit debris are as follows: adjust the two masses on the delivery link of the space target or off-orbit debris.
  • the position of the block adjusts the center of mass of the projecting link to the main connecting shaft, and the moment of inertia of the projecting link rotating around the main connecting shaft is equal to the instantaneous moment of inertia of the space target or off-orbit debris.
  • step S4 the specific adjustment steps of the delivery link of the space target or the off-orbit debris after delivery are as follows:
  • the linear telescoping device performs telescopic operations to adjust the center of mass of the spacecraft system to the delivery plane where the delivery link of the second space target or off-orbit debris to be delivered is rotated vertically around the main connecting axis; when the delivery is reached Upon request, the No. 2 space target or de-orbit debris will be delivered;
  • the present invention is further configured as follows: the specific step of energy dissipation and unloading in step S5 is to use the magnetic moment generated by the orthogonal strong magnetic moment generating device to reversely act on the moment of inertia of the continuously rotating delivery link.
  • the magnetic moment generated by the orthogonal strong magnetic moment generator acts inversely on the moment of inertia of the continuously rotating projection link. In this way, the moment of inertia of the delivery link that rotates vertically around the main connecting shaft is unloaded and energy dissipated until the rotation stops.
  • the present invention is further configured as follows: the specific step of energy dissipating and unloading in step S5 is to use the magnetic moment generated by the orthogonal strong magnetic moment generating device to act in reverse on the residual moment of inertia of the two continuously oppositely rotating projection links.
  • the magnetic moment generated by the orthogonal strong magnetic moment generator acts in the opposite direction on the two continuously rotating projection linkages.
  • the remaining moment of inertia so as to dissipate and unload the moment of inertia of the two delivery links rotating vertically around the main connecting shaft until the rotation stops.
  • Figure 1 is a control flow chart of the suppression method of the present invention
  • FIG. 2 is a schematic diagram of the structure of the spacecraft system according to the first embodiment of the present invention.
  • FIG. 3 is a schematic diagram of the structure of the spacecraft system according to the second embodiment of the present invention.
  • the spacecraft system is divided into a spacecraft main body 1 and a spacecraft main body 2 fixedly connected by a main connecting shaft 3.
  • a projection link 7 perpendicular to the main connecting shaft 3 is rotated on the main connecting shaft 3.
  • the rod 7 is slidably connected with two masses 72 along the length direction, and the center of mass of the spacecraft system is adjusted through the main connecting shaft 3; the specific adjustment steps are as follows:
  • the linear telescopic device connected to the main connecting shaft 3 performs the telescopic operation, extending or contracting the relative position between the spacecraft body 1 and the spacecraft body 2 connected at the two ends of the main connecting shaft 3, measured in the straight line
  • the attitude rotation changes of the spacecraft system are changes in pitch, yaw or roll angle.
  • the center of mass of the spacecraft system does not pass.
  • step (3) Adjust the mass distribution inside the spacecraft body 1 and the spacecraft body 2, and repeat step (2) until the spacecraft system no longer changes in attitude rotation during the expansion and contraction of the linear telescopic device, and the adjustment is completed
  • the center of mass of the spacecraft system passes through the main connecting shaft 3;
  • the linear telescopic device performs telescopic operations until step (4) is repeated in a certain telescopic state of the linear telescopic device, and the spacecraft system no longer changes in attitude rotation, then the center of mass of the spacecraft system is located on the main connecting axis 3 at the same time , In the projection plane where the delivery link 7 rotates vertically around the main connection shaft 3, slide the two masses 72 on the delivery link 7 back to the main connection shaft 3; that is, under no-load conditions, calibrate The telescopic state of the linear telescopic device is the empty 0 position of the corresponding delivery link 7;
  • the adjusted main connecting shaft 3 around the delivery link 7 passes through the center of mass of the spacecraft system after grabbing the space target or de-orbiting debris, and is connected to the delivery
  • the spindles of inertia in the rotation direction of the rod 7 overlap, and the projection plane of the projection link 7 that rotates vertically around the main connecting axis 3 passes through the center of mass of the spacecraft system after grabbing the space target or de-orbiting debris; the specific adjustment steps are as follows:
  • the spacecraft system grabs the space target or off-orbit debris to be projected, holds the space target or off-orbit debris in the corresponding position of the projection link 7, and slides two along the length of the projection link 7.
  • the position of the mass 72 until the measurement spacecraft system no longer undergoes posture rotation changes during the expansion and contraction of the linear telescopic device, that is, the two masses 72 are located on the delivery link 7 of the space target to be delivered or the off-orbit debris.
  • the mass center of the delivery link 7 of the space target or off-orbit debris to be delivered passes through the main connecting shaft 3;
  • the optimized position determined by the two masses 72 of the space target or de-orbit fragments to be delivered is calibrated as the spacecraft system grabbing After the space target or the de-orbited debris, the position 0 of the delivery link 7 to be delivered is completed, that is, the center of mass adjustment after the spacecraft system grabs the space target or de-orbited debris;
  • the delivery link 7 of the space target or off-orbit debris to be delivered is accelerated by energy storage, and drives it to rotate vertically around the main connecting shaft 3; the delivery link 7 adopts the geomagnetic energy storage method for energy storage Speed up rotation
  • the unloading process is the reverse process of energy storage and delivery
  • the energy dissipating unloading is the moment of inertia of the delivery link 7 that rotates vertically around the main connecting shaft 3
  • the magnetic moment generated by the orthogonal strong magnetic moment generator 5 is used Reversely act on the moment of inertia of the continuously rotating delivery link 7 until it stops rotating;
  • the spacecraft system is ready to grab the next space target or off-orbit debris and enter the next delivery cycle.
  • the spacecraft system includes a spacecraft main body 1, a spacecraft main body 2, and the spacecraft main body 1 and a spacecraft main body 2 are fixedly connected by a main connecting shaft 3, and the spacecraft main body system is distributed in the main connection
  • the spacecraft main body 1 and the spacecraft main body 2 at the two ends of the shaft 3 are connected to the main connecting shaft 3 with a linear telescopic mechanism 4 for adjusting the center of mass of the spacecraft system through the main connecting shaft 3, and a rotating mechanism is provided in the middle of the main connecting shaft 3.
  • the projection link 7 perpendicular to the projection link 7 is slidably connected with two masses 72 along the length direction, and the end of the projection link 7 is provided with a holding mechanism 71 for holding space targets or off-track debris.
  • the main connecting shaft 3 is fixedly installed with an orthogonal strong magnetic moment generating device 5, a torque transmission mechanism 6, and the torque transmission mechanism 6 includes a transmission support 61 fixed on the main connecting shaft 3, which is used to drive the delivery link 7
  • the one-way rotating member 62 that rotates around the main connecting shaft 3, the one-way rotating member 62 is rotatably mounted on the transmission support 61; the transmission support 61 of the torque transmission mechanism 6 that starts the work interacts with the one-way rotating member 62
  • the transmission support 61 of the torque transmission mechanism 6 is reacted by the internal torque of the unidirectional rotating member 62 and the orthogonal strong magnetic moment is generated
  • the external torque of the device 5 received by the earth's magnetic field is opposite in direction and of the same magnitude;
  • the orthogonal strong magnetic moment generating device 5 is composed of two orthogonally arranged solenoid coils, and the planes of the two solenoid coils are perpendicular to
  • the main system of the spacecraft includes energy subsystem, control system, communication system, orbit/attitude measurement sensor, solar cell array and auxiliary work load; linear telescopic mechanism 4 is electrically connected to the energy subsystem or solar cell array, and is connected to the controller Phase control connection; the two telescopic ends of the linear telescopic mechanism 4 are respectively fixedly connected to the left half and right half of the main connecting shaft 3. During the telescopic process of the linear telescopic mechanism 4, the left half of the main connecting shaft 3, The right half is on the same line.
  • the orthogonal strong magnetic moment generating device 5 is electrically connected to the energy subsystem or the solar cell array, and is connected to the controller for control;
  • the torque transmission mechanism 6 is electrically connected to the energy subsystem or the solar cell array, and is controlled by the controller
  • the torque transmission mechanism 6 is a torque motor
  • the transmission support 61 is a stator assembly of the torque motor
  • the unidirectional rotating member 62 is a rotor assembly of the torque motor.
  • a method for restraining the nutation of a spacecraft for on-orbit delivery of low-orbit geomagnetic energy storage is disclosed in the present invention, which includes the following control steps:
  • the spacecraft system is divided into a spacecraft body 1 and a spacecraft body 2 which are fixedly connected by a main connecting shaft 3, and two projecting links 7 perpendicular to the main connecting shaft 3 are rotated on the main connecting shaft 3.
  • the connecting rod 7 is slidably connected with two masses 72 along the length direction, and the center of mass of the spacecraft system is adjusted through the main connecting shaft 3; the specific adjustment steps are as follows:
  • the linear telescopic device connected to the main connecting shaft 3 performs the telescopic operation, extending or contracting the relative position between the spacecraft body 1 and the spacecraft body 2 connected at the two ends of the main connecting shaft 3, measured in the straight line
  • the attitude rotation changes of the spacecraft system are changes in pitch, yaw or roll angle.
  • the center of mass of the spacecraft system does not pass.
  • step (3) Adjust the mass distribution inside the spacecraft body 1 and the spacecraft body 2, and repeat step (2) until the spacecraft system no longer changes in attitude rotation during the expansion and contraction of the linear telescopic device, and the adjustment is completed
  • the center of mass of the spacecraft system passes through the main connecting shaft 3;
  • the linear telescopic device performs telescopic operations until step (4) is repeated in a certain telescopic state of the linear telescopic device, and the spacecraft system no longer changes in attitude rotation, then the center of mass of the spacecraft system is located on the main connecting axis 3 at the same time , In the projection plane where the delivery link 7 rotates vertically around the main connection shaft 3, slide the two masses 72 on the delivery link 7 back to the main connection shaft 3; that is, under no-load conditions, calibrate The telescopic state of the linear telescopic device is the empty 0 position of the corresponding delivery link 7;
  • the adjusted main connecting shaft 3 around the delivery link 7 passes through the center of mass of the spacecraft system after grabbing the space target or de-orbiting debris, and is connected to the delivery
  • the spindles of inertia in the rotation direction of the rod 7 overlap, and the projection plane of the projection link 7 that rotates vertically around the main connecting axis 3 passes through the center of mass of the spacecraft system after grabbing the space target or de-orbiting debris; the specific adjustment steps are as follows:
  • the spacecraft system grabs the space target or off-orbit debris to be projected, holds the space target or off-orbit debris in the corresponding position of the projection link 7, and slides two along the length of the projection link 7.
  • the position of the mass 72 until the measurement spacecraft system no longer undergoes posture rotation changes during the expansion and contraction of the linear telescopic device, that is, the two masses 72 are located on the delivery link 7 of the space target to be delivered or the off-orbit debris.
  • the mass center of the delivery link 7 of the space target or off-orbit debris to be delivered passes through the main connecting shaft 3;
  • the delivery link 7 of the space target or off-orbit fragments to be delivered is accelerated by energy storage, and drives it to rotate vertically around the main connecting shaft 3;
  • the linear telescoping device performs telescoping operations to adjust the center of mass of the spacecraft system to the projection plane where the projection link 7 of the space target to be projected or the off-orbit debris rotates vertically around the main connecting axis 3; when it reaches When delivery is required, the No. 2 space target or off-orbit debris will be delivered;
  • the unloading process is the reverse process of energy storage and delivery
  • the energy dissipating unloading is the moment of inertia of the delivery link 7 that rotates vertically around the main connecting shaft 3
  • the magnetic moment generated by the orthogonal strong magnetic moment generator 5 is used Reversely act on the remaining moments of inertia of the two continuously oppositely rotating delivery links 7 until the rotation stops;
  • the spacecraft system is ready to grab the next space target or off-orbit debris and enter the next delivery cycle.
  • the spacecraft system includes a spacecraft main body 1, a spacecraft main body 2, and the spacecraft main body 1 and a spacecraft main body 2 are fixedly connected by a main connecting shaft 3, and the spacecraft main body system is distributed in the main connection
  • the main connecting shaft 3 is connected with a linear telescopic mechanism 4 for adjusting the center of mass of the spacecraft system through the main connecting shaft 3, and the central part of the main connecting shaft 3 rotates with two A projecting link 7 perpendicular to it.
  • the two projecting links 7 are slidably connected with two masses 72 along the length direction, and the ends of the projecting links 7 are provided for keeping the space target or off-track.
  • the counter-rotating transmission mechanism 8 is located between the two projecting links 7, and the counter-rotating transmission mechanism 8 It includes a fixed support 81 fixed on the main connecting shaft 3, a forward rotating member 82 for driving one of the delivery links 7 to rotate in the forward direction around the main connecting shaft 3, and a forward rotation member 82 for driving the other delivery link 7
  • the counter-rotating member 83 that rotates in the opposite direction around the main connecting shaft 3, the forward-rotating member 82 and the counter-rotating member 83 are both rotatably mounted on the fixed support 81; the fixed support 81 of the counter-rotating transmission mechanism 8 that starts the work is respectively
  • the internal torque that interacts with the forward rotating member 82 and the reverse rotating member 83 is formed.
  • the counter-rotating transmission mechanism 8 When the two projection links 7 of the counter-rotating transmission mechanism 8 are in the counter-rotating and rotating projection state, the counter-rotating transmission mechanism 8 is fixed
  • the internal torque of the support 81 reacted by the forward rotating member 82 and the internal torque reacted by the reverse rotating member 83 are opposite in direction and of the same magnitude;
  • the orthogonal strong magnetic moment generating device 5 is composed of two orthogonally arranged solenoid coils,
  • the planes of the two spiral coils are perpendicular to the main connecting axis 3, the orthogonal strong magnetic moment generating device 5 also includes a low temperature system, and the two spiral coils in orthogonal configuration are made of superconductor materials.
  • the main system of the spacecraft includes the energy subsystem, control system, communication system, orbit/attitude measurement sensor, solar cell array and auxiliary work load;
  • the linear telescopic mechanism 4 is electrically connected to the energy subsystem or solar cell array, and is connected to the controller Phase control connection;
  • the two telescopic ends of the linear telescopic mechanism 4 are respectively fixedly connected to the left half and right half of the main connecting shaft 3.
  • the left half of the main connecting shaft 3 The right half is on the same line.
  • the orthogonal strong magnetic moment generating device 5 is electrically connected to the energy subsystem or the solar cell array, and is connected to the controller for control;
  • the counter-rotating transmission mechanism 8 is electrically connected to the energy subsystem or the solar cell array, and is connected to the controller.
  • Control connection, the counter-rotating transmission mechanism 8 is a dual-rotor torque motor, the fixed support 81 is the stator assembly of the dual-rotor torque motor, and the forward rotating member 82 and the reverse rotating member 83 are the two rotating directions of the dual-rotor torque motor respectively.
  • the rotor assembly, and the forward rotating member 82 and the reverse rotating member 83 are coaxially arranged.
  • the spacecraft system adopts a split structure, and the main connecting shaft 3 is connected with a linear telescopic mechanism 4 for measuring whether the center of mass of the spacecraft system passes through the main connecting shaft 3, and the projection link 7 is along the length A mass 72 for adjusting its center of mass is slidably connected in the direction, and the end of the delivery link 7 is provided with a holder for holding space targets or off-orbit debris.
  • the suppression method of this embodiment reasonably allocates the spacecraft attitude rotation change measurement And the sequence of the adjustment of the center of mass makes the spacecraft system's center of mass always be on the main connecting axis 3 and the abrupt moment of inertia before and after the sudden changes in several states during the process of geomagnetic energy storage rotating projection-energy dissipating unloading-re-projection preparation.
  • the spacecraft system adopts a split structure, and the main connecting shaft is connected with a linear telescoping mechanism for measuring whether the spacecraft system's center of mass passes through the main connecting shaft.
  • the projection link is slidably connected in the length direction to adjust its center of mass.
  • the mass block, the end of the delivery link is provided with a cage for holding space targets or off-orbit debris.
  • the suppression method of the present invention reasonably allocates the sequence of attitude rotation measurement and center of mass adjustment, so that the geomagnetic energy storage rotating delivery-dissipation Can be unloaded-before and after the sudden change of several states in the process of re-projection preparation, the center of mass of the spacecraft system is always at the intersection of the main connecting axis and the sudden moment of inertia plane, thereby effectively preventing the spacecraft system from failing due to the rotating inertia main axis
  • the center of mass causes nutation problems;
  • the magnetic moment generated by the orthogonal strong magnetic moment generator acts in the opposite direction on the moment of inertia of the continuously rotating delivery link, thereby dissipating energy and unloading the delivery link that rotates vertically around the main connecting shaft.
  • the main connecting shaft around which the delivery link of the space target or the de-orbiting debris is realized passes through the spacecraft after grabbing the space target or the de-orbiting debris.
  • the center of mass of the system is overlapped with the main axis of inertia in the direction of rotation of the projection link.
  • the projection plane of the projection link that rotates vertically around the main link axis passes through the spacecraft system after grabbing space targets or off-orbit debris. Centroid.

Abstract

一种低轨道地磁蓄能在轨投送的航天器章动抑制方法，包括：S1，投送连杆（7）沿长度方向上滑动连接有两个质量块（72），调整航天器系统的质心通过主连接轴（3）；S2，在航天器系统抓取空间目标或离轨碎片，且空间目标或离轨碎片保持于投送连杆（7）的相应位置后，对待投送空间目标或离轨碎片的投送连杆（7）的质心和惯量主轴分别进行测量标定和调整；S3，蓄能投送；S4，对投送完空间目标或离轨碎片的投送连杆（7）的质心和转动惯量分别进行调整；S5，消能卸载；S6，航天器系统准备抓取下一个空间目标或离轨碎片，进入下一个投送工作循环。

Description

一种低轨道地磁蓄能在轨投送的航天器章动抑制方法 技术领域
本发明涉及低轨道地磁蓄能在轨投送技术领域。
背景技术
航天器系统在低轨道空间使用地磁蓄能方法进行在轨投送目标时，会存在一个重大问题是：航天器系统的投送机构所旋绕的旋转轴可能不通过航天器系统的质心，这会导致航天器系统发生旋转章动现象，在低轨道空间中没有空气阻尼的环境下，章动现象对于航天器系统是非常危险的。
影响航天器系统的旋转轴不通过其质心有以下几个方面的因素：1)航天器系统的自身状态发生改变，比如在轨工作时消耗了部分燃料，或者是其所携带的设备、负载等发生了位置移动或旋动；2)航天器系统抓取了不确定的空间目标或者离轨碎片，导致航天器系统的质量和质心均发生不确定变化；3)航天器系统投送空间目标或者离轨碎片的瞬时，航天器系统的质量和质心亦会发生变化。
针对航天器系统在投送空间目标或者离轨碎片的瞬时，航天器系统的质量和质心也会发生相应变化的情况，亟需设计研发一种低轨道地磁蓄能在轨投送的航天器章动抑制方法。
发明内容
本发明的目的是提供一种低轨道地磁蓄能在轨投送的航天器章动抑制方法，在地磁蓄能旋转投送-消能卸载-再次投送准备过程中的几个状态突变前后，有效地抑制航天器系统在轨自由章动的能力。
本发明的上述目的是通过以下技术方案得以实现的：
一种低轨道地磁蓄能在轨投送的航天器章动抑制方法，包括有以下控制步骤：
S1，航天器系统分为由主连接轴相固接的航天器主体一、航天器主体二，主连接轴上转动设置有与其相垂直的投送连杆，投送连杆沿长度方向上滑动连接有两个质量块，调整航天器系统的质心通过主连接轴；
S2，在航天器系统抓取空间目标或离轨碎片后，空间目标或离轨碎片保持于投送连杆的相应位置后，对待投送空间目标或离轨碎片的投送连杆的质心和惯量主轴分别进行测量标定和调整，调整后的该投送连杆所绕的主连接轴通过抓取空间目标或离轨碎片后的航天器系统的质心，且与该投送连杆旋转方向的惯量主轴相重叠，同时该投送连杆绕主连接轴垂直旋转的投送平面通过抓取空间目标或离轨碎片后的航天器系统的质心；
S3，蓄能投送：对待投送空间目标或离轨碎片的投送连杆进行蓄能加速，驱动其绕主连接轴垂直旋转；
S4，当达到空间目标或离轨碎片的投送要求时，即投送空间目标或离轨碎片，对投送完空间目标或离轨碎片的投送连杆的质心和转动惯量分别进行调整；
S5，消能卸载：卸载过程为蓄能投送的逆过程，消能卸载绕主连接轴垂直旋转的投送连杆的转动惯量，直至停止旋转；
S6，航天器系统准备抓取下一个空间目标或离轨碎片，进入下一个投送工作循环。
通过采用上述技术方案，航天器系统采用分体式结构，并在主连接轴上联接有用于测量航天器系统质心是否通过主连接轴的直线伸缩机构，投送连杆沿长度方向上滑动连接有用于调节其质心的质量块，投送连杆的端部设有用于保持空间目标或离轨碎片的保持架，并合理分配姿态旋转测量和质心调整的顺序，使得在地磁蓄能旋转投送-消能卸载-再次投送准备的过程中几个状态的突变前后，航天器系统的质心始终处于主连接轴与突变的转动惯量平面内的交点，从而有效抑制航天器系统由于旋转的惯量主轴不通过质心而造 成章动问题。
本发明进一步设置为：所述步骤S1的具体调整步骤如下：
(1)在航天器系统抓取待投送的空间目标或离轨碎片之前，将投送连杆上的两个质量块均滑动回至主连接轴上；
(2)联接于主连接轴上的直线伸缩装置执行伸缩作业，伸长或收缩主连接轴两端连接的航天器主体一、航天器主体二之间的相对位置，测量在直线伸缩装置的伸缩过程中航天器系统的姿态旋转变化，当航天器系统发生姿态旋转变化，即得航天器系统的质心不通过主连接轴；
(3)调整航天器主体一、航天器主体二内部的质量分布，再重复步骤(2)，直至在直线伸缩装置的伸缩过程中，航天器系统不再发生姿态旋转变化，即完成调整航天器系统的质心通过主连接轴；
(4)滑动投送连杆上两个质量块的位置，在质量块滑动过程中，测量航天器系统发生姿态旋转变化，即得航天器系统的质心不通过投送连杆绕主连接轴垂直旋转的投送平面内；
(5)直线伸缩装置执行伸缩作业，直至在直线伸缩装置的某一伸缩状态下重复步骤(4)，航天器系统不再发生姿态旋转变化，则航天器系统的质心同时位于主连接轴上、投送连杆绕主连接轴垂直旋转的投送平面内，再将投送连杆上的两个质量块滑动回至主连接轴上；即在空载条件下，标定直线伸缩装置的该伸缩状态为相对应的投送连杆的空载0位；
(6)按照上述步骤分别标定在空载条件下各个投送连杆相对应的直线伸缩装置的伸缩状态的空载0位。
通过采用上述技术方案，在航天器系统抓取空间目标或离轨碎片之前，沿投送连杆长度方向上滑动两个质量块，调整航天器系统的质心通过主连接轴，从而标定在空载条件下各个投送连杆相对应的直线伸缩装置的伸缩状态的空载0位。
本发明进一步设置为：所述步骤S2的具体调整步骤如下：
(1)航天器系统抓取待投送的空间目标或离轨碎片，将空间目标或离轨碎片保持于投送连杆的相应位置后，沿投送连杆的长度方向滑动两个质量块的位置，直至在直线伸缩装置的伸缩过程中，测量航天器系统不再发生姿态旋转变化，即为两个质量块位于待投送空间目标或离轨碎片的投送连杆上的平衡位置，待投送空间目标或离轨碎片的投送连杆的质心通过主连接轴；
(2)直线伸缩装置调整至与该投送连杆相对应的空载0位的伸缩状态；
(3)分析该投送连杆上质量确定的两个质量块的位置，使得两个质量块相对于其所绕的主连接轴的转动惯量最小，并结合两个质量块位于待投送空间目标或离轨碎片的投送连杆上的平衡位置，即得待投送空间目标或离轨碎片的两个质量块所定的优化位置，标定为航天器系统抓取空间目标或离轨碎片后该投送连杆的待投送0位；
(4)按照上述步骤分别标定航天器系统抓取空间目标或离轨碎片后各个投送连杆的待投送0位，即完成航天器系统抓取空间目标或离轨碎片后的质心调整。
通过采用上述技术方案，对待投送空间目标或离轨碎片的投送连杆的质心和惯量主轴分别进行测量标定和调整，实现该投送连杆所绕的主连接轴通过抓取空间目标或离轨碎片后的航天器系统的质心，且与该投送连杆旋转方向的惯量主轴相重叠，同时该投送连杆绕主连接轴垂直旋转的投送平面通过抓取空间目标或离轨碎片后的航天器系统的质心。
本发明进一步设置为：所述航天器系统发生的姿态旋转变化为俯仰、偏航或滚转角度变化。
通过采用上述技术方案，通过测量航天器系统发生的姿态旋转变化，实现判断航天器系统的质心是否通过主连接轴的目的。
本发明进一步设置为：当所述主连接轴上转动设置有一个与其相垂直的 投送连杆，所述步骤S3中投送连杆采用地磁蓄能方法进行蓄能加速旋转。
通过采用上述技术方案，针对主连接轴上转动设置有一个投送连杆的航天器系统，正交强磁矩生成装置在地磁场的外力矩与反作用于力矩传动机构的传动支座受到单向旋转件的内力矩相平衡，航天器系统在蓄能加速过程中不会发生姿态的加速旋转的情况。
本发明进一步设置为：当所述主连接轴上转动设置有两个与其相垂直的投送连杆，所述步骤S3中两个投送连杆采用对转传动机构驱动其进行蓄能加速反向旋转。
通过采用上述技术方案，针对主连接轴上转动设置有两个投送连杆的航天器系统，不采用地磁蓄能的加速方式，对转传动机构的固定支座受到正向旋转件反作用的内力矩与受到反向旋转件反作用的内力矩相平衡，航天器系统在蓄能加速过程中不会发生姿态的加速旋转的情况。
本发明进一步设置为：所述步骤S4中对投送完空间目标或离轨碎片的投送连杆具体调整步骤如下：调整投送完空间目标或离轨碎片的投送连杆上两个质量块的位置，将该投送连杆的质心调回至主连接轴上，且该投送连杆绕主连接轴旋转的转动惯量与投送完空间目标或离轨碎片的瞬时转动惯量相等。
通过采用上述技术方案，针对主连接轴上转动设置有一个投送连杆的航天器系统，完成投送完空间目标或离轨碎片的投送连杆的质心和转动惯量的调整作业。
本发明进一步设置为：所述步骤S4中对投送完空间目标或离轨碎片的投送连杆具体调整步骤如下：
(1)调整投送完一号空间目标或离轨碎片的投送连杆上两个质量块的位置，将该投送连杆的质心调回至主连接轴上，且该投送连杆绕主连接轴旋转的转动惯量与投送完一号空间目标或离轨碎片的瞬时转动惯量相等；
(2)直线伸缩装置执行伸缩作业，将航天器系统的质心调整至待投送二号空间目标或离轨碎片的投送连杆绕主连接轴垂直旋转的投送平面内；当达到投送要求时，即投送二号空间目标或离轨碎片；
(3)调整投送完二号空间目标或离轨碎片的投送连杆上两个质量块的位置，将该投送连杆的质心调回至主连接轴上，且该投送连杆绕主连接轴旋转的转动惯量与投送完二号空间目标或离轨碎片的瞬时转动惯量相等。
通过采用上述技术方案，针对主连接轴上转动设置有两个投送连杆的航天器系统，依次完成投送完空间目标或离轨碎片的两个投送连杆的质心和转动惯量的调整作业。
本发明进一步设置为：所述步骤S5中消能卸载的具体步骤为采用正交强磁矩生成装置产生的磁力矩反向作用于持续旋转的投送连杆的转动惯量。
通过采用上述技术方案，针对主连接轴上转动设置有一个投送连杆的航天器系统，正交强磁矩生成装置产生的磁力矩反向作用于持续旋转的投送连杆的转动惯量，从而消能卸载绕主连接轴垂直旋转的投送连杆的转动惯量，直至停止旋转。
本发明进一步设置为：所述步骤S5中消能卸载的具体步骤为采用正交强磁矩生成装置产生的磁力矩反向作用于两个持续对向旋转的投送连杆的剩余转动惯量。
通过采用上述技术方案，针对主连接轴上转动设置有两个投送连杆的航天器系统，正交强磁矩生成装置产生的磁力矩反向作用于持续旋转的两个投送连杆的剩余转动惯量，从而消能卸载绕主连接轴垂直旋转的两个投送连杆的转动惯量，直至停止旋转。
附图说明
图1是本发明抑制方法的控制流程图；
图2是本发明实施例一的航天器系统结构示意图；
图3是本发明实施例二的航天器系统结构示意图；
图中标号，1、航天器主体一；2、航天器主体二；3、主连接轴；4、直线伸缩机构；5、正交强磁矩生成装置；6、力矩传动机构；61、传动支座；62、单向旋转件；7、投送连杆；71、保持机构；72、质量块；8、对转传动机构；81、固定支座；82、正向旋转件；83、反向旋转件。
本发明的较佳实施方式
参照图1、2，为本发明公开的一种低轨道地磁蓄能在轨投送的航天器章动抑制方法，包括有以下控制步骤：
S1，航天器系统分为由主连接轴3相固接的航天器主体一1、航天器主体二2，主连接轴3上转动设置有一个与其相垂直的投送连杆7，投送连杆7沿长度方向上滑动连接有两个质量块72，调整航天器系统的质心通过主连接轴3；具体调整步骤如下：
(1)在航天器系统抓取待投送的空间目标或离轨碎片之前，将投送连杆7上的两个质量块72均滑动回至主连接轴3上；
(2)联接于主连接轴3上的直线伸缩装置执行伸缩作业，伸长或收缩主连接轴3两端连接的航天器主体一1、航天器主体二2之间的相对位置，测量在直线伸缩装置的伸缩过程中航天器系统姿态旋转变化，航天器系统发生的姿态旋转变化为俯仰、偏航或滚转角度变化，当航天器系统发生姿态旋转变化，即得航天器系统的质心不通过主连接轴3；
(3)调整航天器主体一1、航天器主体二2内部的质量分布，再重复步骤(2)，直至在直线伸缩装置的伸缩过程中，航天器系统不再发生姿态旋转变化，即完成调整航天器系统的质心通过主连接轴3；
(4)滑动投送连杆7上两个质量块72的位置，在质量块72滑动过程中，测量航天器系统发生姿态旋转变化，即得航天器系统的质心不通过投送连杆7绕主连接轴3垂直旋转的投送平面内；
(5)直线伸缩装置执行伸缩作业，直至在直线伸缩装置的某一伸缩状态 下重复步骤(4)，航天器系统不再发生姿态旋转变化，则航天器系统的质心同时位于主连接轴3上、投送连杆7绕主连接轴3垂直旋转的投送平面内，再将投送连杆7上的两个质量块72滑动回至主连接轴3上；即在空载条件下，标定直线伸缩装置的该伸缩状态为相对应的投送连杆7的空载0位；
S2，在航天器系统抓取空间目标或离轨碎片后，空间目标或离轨碎片保持于投送连杆7的相应位置后，对待投送空间目标或离轨碎片的投送连杆7的质心和惯量主轴分别进行测量标定和调整，调整后的该投送连杆7所绕的主连接轴3通过抓取空间目标或离轨碎片后的航天器系统的质心，且与该投送连杆7旋转方向的惯量主轴相重叠，同时该投送连杆7绕主连接轴3垂直旋转的投送平面通过抓取空间目标或离轨碎片后的航天器系统的质心；具体调整步骤如下：
(1)航天器系统抓取待投送的空间目标或离轨碎片，将空间目标或离轨碎片保持于投送连杆7的相应位置后，沿投送连杆7的长度方向滑动两个质量块72的位置，直至在直线伸缩装置的伸缩过程中，测量航天器系统不再发生姿态旋转变化，即为两个质量块72位于待投送空间目标或离轨碎片的投送连杆7上的平衡位置，待投送空间目标或离轨碎片的投送连杆7的质心通过主连接轴3；
(2)直线伸缩装置调整至与该投送连杆7相对应的空载0位的伸缩状态；
(3)分析该投送连杆7上质量确定的两个质量块72的位置，使得两个质量块72相对于其所绕的主连接轴3的转动惯量最小，并结合两个质量块72位于待投送空间目标或离轨碎片的投送连杆7上的平衡位置，即得待投送空间目标或离轨碎片的两个质量块72所定的优化位置，标定为航天器系统抓取空间目标或离轨碎片后该投送连杆7的待投送0位，即完成航天器系统抓取空间目标或离轨碎片后的质心调整；
S3，蓄能投送：对待投送空间目标或离轨碎片的投送连杆7进行蓄能加 速，驱动其绕主连接轴3垂直旋转；投送连杆7采用地磁蓄能方法进行蓄能加速旋转；
S4，当达到空间目标或离轨碎片的投送要求时，即投送空间目标或离轨碎片；对投送完空间目标或离轨碎片的投送连杆7的质心和转动惯量分别进行调整；具体调整步骤如下：调整投送完空间目标或离轨碎片的投送连杆7上两个质量块72的位置，将该投送连杆7的质心调回至主连接轴3上，且该投送连杆7绕主连接轴3旋转的转动惯量与投送完空间目标或离轨碎片的瞬时转动惯量相等；
S5，消能卸载：卸载过程为蓄能投送的逆过程，消能卸载绕主连接轴3垂直旋转的投送连杆7的转动惯量，采用正交强磁矩生成装置5产生的磁力矩反向作用于持续旋转的投送连杆7的转动惯量，直至停止旋转；
S6，航天器系统准备抓取下一个空间目标或离轨碎片，进入下一个投送工作循环。
其中，该航天器系统包括有航天器主体一1、航天器主体二2，航天器主体一1与航天器主体二2之间通过主连接轴3相固定连接，航天器主体系统分布于主连接轴3两端的航天器主体一1、航天器主体二2上；主连接轴3上联接有用于调节航天器系统质心通过主连接轴3的直线伸缩机构4，主连接轴3中部转动设置有一个与其相垂直的投送连杆7，投送连杆7沿长度方向上滑动连接有两个质量块72，且投送连杆7端部设有用于保持空间目标或离轨碎片的保持机构71；主连接轴3上固定安装有正交强磁矩生成装置5、力矩传动机构6，力矩传动机构6包括有固定于主连接轴3上的传动支座61、用于驱动投送连杆7绕主连接轴3旋转的单向旋转件62，单向旋转件62转动安装于传动支座61上；启动工作的力矩传动机构6的传动支座61与单向旋转件62之间形成相互作用的内力矩，当力矩传动机构6的投送连杆7处于地磁蓄能旋转投送状态，力矩传动机构6的传动支座61受到单向旋转件62 反作用的内力矩与正交强磁矩生成装置5受到地磁场的外力矩为方向相反、大小相同；正交强磁矩生成装置5由两个正交配置的螺线圈构成，且两个螺线圈的平面均与主连接轴3相垂直，正交强磁矩生成装置5还包括有低温系统，且两个正交配置的螺线圈均采用超导体材料制成。
航天器主体系统包括能源子系统、控制系统、通讯系统、轨/姿测量敏感器、太阳能电池阵列及辅助作业载荷；直线伸缩机构4与能源子系统或太阳能电池阵列相电连接，并与控制器相控制连接；直线伸缩机构4的两个伸缩端分别与主连接轴3的左半段、右半段相固接，在直线伸缩机构4的伸缩过程中，主连接轴3的左半段、右半段位于同一直线上。正交强磁矩生成装置5与能源子系统或太阳能电池阵列相电连接，并与控制器相控制连接；力矩传动机构6与能源子系统或太阳能电池阵列相电连接，并与控制器相控制连接，力矩传动机构6为力矩电机，传动支座61为力矩电机的定子组件，单向旋转件62为力矩电机的转子组件。
实施例二：
参见图1、3，为本发明公开的一种低轨道地磁蓄能在轨投送的航天器章动抑制方法，包括有以下控制步骤：
S1，航天器系统分为由主连接轴3相固接的航天器主体一1、航天器主体二2，主连接轴3上转动设置有两个与其相垂直的投送连杆7，投送连杆7沿长度方向上滑动连接有两个质量块72，调整航天器系统的质心通过主连接轴3；具体调整步骤如下：
(1)在航天器系统抓取待投送的空间目标或离轨碎片之前，将投送连杆7上的两个质量块72均滑动回至主连接轴3上；
(2)联接于主连接轴3上的直线伸缩装置执行伸缩作业，伸长或收缩主连接轴3两端连接的航天器主体一1、航天器主体二2之间的相对位置，测量在直线伸缩装置的伸缩过程中航天器系统姿态旋转变化，航天器系统发生的 姿态旋转变化为俯仰、偏航或滚转角度变化，当航天器系统发生姿态旋转变化，即得航天器系统的质心不通过主连接轴3；
(3)调整航天器主体一1、航天器主体二2内部的质量分布，再重复步骤(2)，直至在直线伸缩装置的伸缩过程中，航天器系统不再发生姿态旋转变化，即完成调整航天器系统的质心通过主连接轴3；
(4)滑动投送连杆7上两个质量块72的位置，在质量块72滑动过程中，测量航天器系统发生姿态旋转变化，即得航天器系统的质心不通过投送连杆7绕主连接轴3垂直旋转的投送平面内；
(5)直线伸缩装置执行伸缩作业，直至在直线伸缩装置的某一伸缩状态下重复步骤(4)，航天器系统不再发生姿态旋转变化，则航天器系统的质心同时位于主连接轴3上、投送连杆7绕主连接轴3垂直旋转的投送平面内，再将投送连杆7上的两个质量块72滑动回至主连接轴3上；即在空载条件下，标定直线伸缩装置的该伸缩状态为相对应的投送连杆7的空载0位；
(6)按照上述步骤分别标定在空载条件下各个投送连杆7相对应的直线伸缩装置的伸缩状态的空载0位；
S2，在航天器系统抓取空间目标或离轨碎片后，空间目标或离轨碎片保持于投送连杆7的相应位置后，对待投送空间目标或离轨碎片的投送连杆7的质心和惯量主轴分别进行测量标定和调整，调整后的该投送连杆7所绕的主连接轴3通过抓取空间目标或离轨碎片后的航天器系统的质心，且与该投送连杆7旋转方向的惯量主轴相重叠，同时该投送连杆7绕主连接轴3垂直旋转的投送平面通过抓取空间目标或离轨碎片后的航天器系统的质心；具体调整步骤如下：
(1)航天器系统抓取待投送的空间目标或离轨碎片，将空间目标或离轨碎片保持于投送连杆7的相应位置后，沿投送连杆7的长度方向滑动两个质量块72的位置，直至在直线伸缩装置的伸缩过程中，测量航天器系统不再发 生姿态旋转变化，即为两个质量块72位于待投送空间目标或离轨碎片的投送连杆7上的平衡位置，待投送空间目标或离轨碎片的投送连杆7的质心通过主连接轴3；
(2)直线伸缩装置调整至与该投送连杆7相对应的空载0位的伸缩状态；
(3)分析该投送连杆7上质量确定的两个质量块72的位置，使得两个质量块72相对于其所绕的主连接轴3的转动惯量最小，并结合两个质量块72位于待投送空间目标或离轨碎片的投送连杆7上的平衡位置，即得待投送空间目标或离轨碎片的两个质量块72所定的优化位置，标定为航天器系统抓取空间目标或离轨碎片后该投送连杆7的待投送0位；
(4)按照上述步骤分别标定航天器系统抓取空间目标或离轨碎片后各个投送连杆7的待投送0位；即完成航天器系统抓取空间目标或离轨碎片后的质心调整；
S3，蓄能投送：对待投送空间目标或离轨碎片的投送连杆7进行蓄能加速，驱动其绕主连接轴3垂直旋转；两个投送连杆7采用对转传动机构8驱动其进行蓄能加速反向旋转；
S4，当达到空间目标或离轨碎片的投送要求时，即投送空间目标或离轨碎片；对投送完空间目标或离轨碎片的投送连杆7的质心和转动惯量分别进行调整；具体调整步骤如下：
(1)调整投送完一号空间目标或离轨碎片的投送连杆7上两个质量块72的位置，将该投送连杆7的质心调回至主连接轴3上，且该投送连杆7绕主连接轴3旋转的转动惯量与投送完一号空间目标或离轨碎片的瞬时转动惯量相等；
(2)直线伸缩装置执行伸缩作业，将航天器系统的质心调整至待投送二号空间目标或离轨碎片的投送连杆7绕主连接轴3垂直旋转的投送平面内；当达到投送要求时，即投送二号空间目标或离轨碎片；
(3)调整投送完二号空间目标或离轨碎片的投送连杆7上两个质量块72的位置，将该投送连杆7的质心调回至主连接轴3上，且该投送连杆7绕主连接轴3旋转的转动惯量与投送完二号空间目标或离轨碎片的瞬时转动惯量相等；
S5，消能卸载：卸载过程为蓄能投送的逆过程，消能卸载绕主连接轴3垂直旋转的投送连杆7的转动惯量，采用正交强磁矩生成装置5产生的磁力矩反向作用于两个持续对向旋转的投送连杆7的剩余转动惯量，直至停止旋转；
S6，航天器系统准备抓取下一个空间目标或离轨碎片，进入下一个投送工作循环。
其中，该航天器系统包括有航天器主体一1、航天器主体二2，航天器主体一1与航天器主体二2之间通过主连接轴3相固定连接，航天器主体系统分布于主连接轴3两端的航天器主体一1、航天器主体二2上；主连接轴3上联接有用于调节航天器系统质心通过主连接轴3的直线伸缩机构4，主连接轴3中部转动设置有两个与其相垂直的投送连杆7，两个投送连杆7沿长度方向上均滑动连接有两个质量块72，且投送连杆7端部均设有用于保持空间目标或离轨碎片的保持机构71；主连接轴3上固定安装有正交强磁矩生成装置5、对转传动机构8，对转传动机构8位于两个投送连杆7之间，对转传动机构8包括有固定于主连接轴3上的固定支座81、用于驱动其中一个投送连杆7绕主连接轴3正向旋转的正向旋转件82、用于驱动另一个投送连杆7绕主连接轴3反向旋转的反向旋转件83，正向旋转件82、反向旋转件83均转动安装于固定支座81上；启动工作的对转传动机构8的固定支座81分别与正向旋转件82、反向旋转件83之间形成相互作用的内力矩，当对转传动机构8的两个投送连杆7处于对转旋转投送状态，对转传动机构8的固定支座81受到正向旋转件82反作用的内力矩与受到反向旋转件83反作用的内力 矩为方向相反、大小相同；正交强磁矩生成装置5由两个正交配置的螺线圈构成，且两个螺线圈的平面均与主连接轴3相垂直，正交强磁矩生成装置5还包括有低温系统，且两个正交配置的螺线圈均采用超导体材料制成。
航天器主体系统包括能源子系统、控制系统、通讯系统、轨/姿测量敏感器、太阳能电池阵列及辅助作业载荷；直线伸缩机构4与能源子系统或太阳能电池阵列相电连接，并与控制器相控制连接；直线伸缩机构4的两个伸缩端分别与主连接轴3的左半段、右半段相固接，在直线伸缩机构4的伸缩过程中，主连接轴3的左半段、右半段位于同一直线上。正交强磁矩生成装置5与能源子系统或太阳能电池阵列相电连接，并与控制器相控制连接；对转传动机构8与能源子系统或太阳能电池阵列相电连接，并与控制器相控制连接，对转传动机构8为双转子力矩电机，固定支座81为双转子力矩电机的定子组件，正向旋转件82、反向旋转件83分别为双转子力矩电机的两个旋向相反的转子组件，且正向旋转件82与反向旋转件83为同轴设置。
本实施例的实施原理为：航天器系统采用分体式结构，并在主连接轴3上联接有用于测量航天器系统质心是否通过主连接轴3的直线伸缩机构4，投送连杆7沿长度方向上滑动连接有用于调节其质心的质量块72，投送连杆7的端部设有用于保持空间目标或离轨碎片的保持架，本实施例的抑制方法合理分配航天器姿态旋转变化测量和质心调整的顺序，使得在地磁蓄能旋转投送-消能卸载-再次投送准备的过程中几个状态的突变前后，航天器系统的质心始终处于主连接轴3与突变的转动惯量平面内的交点，再完成投送完空间目标或离轨碎片的投送连杆7的质心和转动惯量的调整作业，并消能卸载绕主连接轴3垂直旋转的投送连杆7的转动惯量，采用正交强磁矩生成装置5产生的磁力矩反向作用于持续旋转的投送连杆7的转动惯量，直至停止旋转。最终实现在地磁蓄能旋转投送-消能卸载-再次投送准备过程中的几个状态突变前后，有效地抑制航天器系统在轨自由章动的能力。
本具体实施方式的实施例均为本发明的较佳实施例，并非依此限制本发明的保护范围，故：凡依本发明的结构、形状、原理所做的等效变化，均应涵盖于本发明的保护范围之内。
工业实用性
1.航天器系统采用分体式结构，并在主连接轴上联接有用于测量航天器系统质心是否通过主连接轴的直线伸缩机构，投送连杆沿长度方向上滑动连接有用于调节其质心的质量块，投送连杆的端部设有用于保持空间目标或离轨碎片的保持架，本发明的抑制方法合理分配姿态旋转测量和质心调整的顺序，使得在地磁蓄能旋转投送-消能卸载-再次投送准备的过程中几个状态的突变前后，航天器系统的质心始终处于主连接轴与突变的转动惯量平面内的交点，从而有效抑制航天器系统由于旋转的惯量主轴不通过质心而造成章动问题；
2.在消能卸载过程中，正交强磁矩生成装置产生的磁力矩反向作用于持续旋转的投送连杆的转动惯量，从而消能卸载绕主连接轴垂直旋转的投送连杆的转动惯量，直至停止旋转，以防止航天器系统发生姿态旋转变化；
3.在航天器系统抓取空间目标或离轨碎片后，实现待投送空间目标或离轨碎片的投送连杆所绕的主连接轴通过抓取空间目标或离轨碎片后的航天器系统的质心，且与该投送连杆旋转方向的惯量主轴相重叠，同时该投送连杆绕主连接轴垂直旋转的投送平面通过抓取空间目标或离轨碎片后的航天器系统的质心。

Claims (10)

  1. 一种低轨道地磁蓄能在轨投送的航天器章动抑制方法，其特征在于，包括有以下控制步骤：
    S1，航天器系统分为由主连接轴(3)相固接的航天器主体一(1)、航天器主体二(2)，主连接轴(3)上转动设置有与其相垂直的投送连杆(7)，投送连杆(7)沿长度方向上滑动连接有两个质量块(72)，调整航天器系统的质心通过主连接轴(3)；
    S2，在航天器系统抓取空间目标或离轨碎片后，空间目标或离轨碎片保持于投送连杆(7)的相应位置后，对待投送空间目标或离轨碎片的投送连杆(7)的质心和惯量主轴分别进行测量标定和调整，调整后的该投送连杆(7)所绕的主连接轴(3)通过抓取空间目标或离轨碎片后的航天器系统的质心，且与该投送连杆(7)旋转方向的惯量主轴相重叠，同时该投送连杆(7)绕主连接轴(3)垂直旋转的投送平面通过抓取空间目标或离轨碎片后的航天器系统的质心；
    S3，蓄能投送：对待投送空间目标或离轨碎片的投送连杆(7)进行蓄能加速，驱动其绕主连接轴(3)垂直旋转；
    S4，当达到空间目标或离轨碎片的投送要求时，即投送空间目标或离轨碎片，对投送完空间目标或离轨碎片的投送连杆(7)的质心和转动惯量分别进行调整；
    S5，消能卸载：卸载过程为蓄能投送的逆过程，消能卸载绕主连接轴(3)垂直旋转的投送连杆(7)的转动惯量，直至停止旋转；
    S6，航天器系统准备抓取下一个空间目标或离轨碎片，进入下一个投送工作循环。
  2. 根据权利要求1所述的航天器章动抑制方法，其特征在于，所述步骤S1的具体调整步骤如下：
    (1)在航天器系统抓取待投送的空间目标或离轨碎片之前，将投送连杆(7)上的两个质量块(72)均滑动回至主连接轴(3)上；
    (2)联接于主连接轴(3)上的直线伸缩装置执行伸缩作业，伸长或收缩主连接轴(3)两端连接的航天器主体一(1)、航天器主体二(2)之间的相对位置，测量在直线伸缩装置的伸缩过程中航天器系统的姿态旋转变化，当航天器系统发生姿态旋转变化，即得航天器系统的质心不通过主连接轴(3)；
    (3)调整航天器主体一(1)、航天器主体二(2)内部的质量分布，再重复步骤(2)，直至在直线伸缩装置的伸缩过程中，航天器系统不再发生姿态旋转变化，即完成调整航天器系统的质心通过主连接轴(3)；
    (4)滑动投送连杆(7)上两个质量块(72)的位置，在质量块(72)滑动过程中，测量航天器系统发生姿态旋转变化，即得航天器系统的质心不通过投送连杆(7)绕主连接轴(3)垂直旋转的投送平面内；
    (5)直线伸缩装置执行伸缩作业，直至在直线伸缩装置的某一伸缩状态下重复步骤(4)，航天器系统不再发生姿态旋转变化，则航天器系统的质心同时位于主连接轴(3)上、投送连杆(7)绕主连接轴(3)垂直旋转的投送平面内，再将投送连杆(7)上的两个质量块(72)滑动回至主连接轴(3)上；即在空载条件下，标定直线伸缩装置的该伸缩状态为相对应的投送连杆(7)的空载0位；
    (6)按照上述步骤分别标定在空载条件下各个投送连杆(7)相对应的直线伸缩装置的伸缩状态的空载0位。
  3. 根据权利要求1所述的航天器章动抑制方法，其特征在于，所述步骤S2的具体调整步骤如下：
    (1)航天器系统抓取待投送的空间目标或离轨碎片，将空间目标或离轨碎片保持于投送连杆(7)的相应位置后，沿投送连杆(7)的长度方向滑动两个质量块(72)的位置，直至在直线伸缩装置的伸缩过程中，测量航天器系统 不再发生姿态旋转变化，即为两个质量块(72)位于待投送空间目标或离轨碎片的投送连杆(7)上的平衡位置，待投送空间目标或离轨碎片的投送连杆(7)的质心通过主连接轴(3)；
    (2)直线伸缩装置调整至与该投送连杆(7)相对应的空载0位的伸缩状态；
    (3)分析该投送连杆(7)上质量确定的两个质量块(72)的位置，使得两个质量块(72)相对于其所绕的主连接轴(3)的转动惯量最小，并结合两个质量块(72)位于待投送空间目标或离轨碎片的投送连杆(7)上的平衡位置，即得待投送空间目标或离轨碎片的两个质量块(72)所定的优化位置，标定为航天器系统抓取空间目标或离轨碎片后该投送连杆(7)的待投送0位；
    (4)按照上述步骤分别标定航天器系统抓取空间目标或离轨碎片后各个投送连杆(7)的待投送0位，即完成航天器系统抓取空间目标或离轨碎片后的质心调整。
  4. 根据权利要求1或2所述的航天器章动抑制方法，其特征在于，所述航天器系统发生的姿态旋转变化为俯仰、偏航或滚转角度变化。
  5. 根据权利要求1～3任一项所述的航天器章动抑制方法，其特征在于，所述主连接轴(3)上转动设置有一个与其相垂直的投送连杆(7)，所述步骤S3中投送连杆(7)采用地磁蓄能方法进行蓄能加速旋转。
  6. 根据权利要求1～3任一项所述的航天器章动抑制方法，其特征在于，所述主连接轴(3)上转动设置有两个与其相垂直的投送连杆(7)，所述步骤S3中两个投送连杆(7)采用对转传动机构(8)驱动其进行蓄能加速反向旋转。
  7. 根据权利要求4所述的航天器章动抑制方法，其特征在于，所述步骤S4中对投送完空间目标或离轨碎片的投送连杆(7)具体调整步骤如下：调整投送完空间目标或离轨碎片的投送连杆(7)上两个质量块(72)的位置，将该投送连杆(7)的质心调回至主连接轴(3)上，且该投送连杆(7)绕主连接轴(3)旋转 的转动惯量与投送完空间目标或离轨碎片的瞬时转动惯量相等。
  8. 根据权利要求5所述的航天器章动抑制方法，其特征在于，所述步骤S4中对投送完空间目标或离轨碎片的投送连杆(7)具体调整步骤如下：
    (1)调整投送完一号空间目标或离轨碎片的投送连杆(7)上两个质量块(72)的位置，将该投送连杆(7)的质心调回至主连接轴(3)上，且该投送连杆(7)绕主连接轴(3)旋转的转动惯量与投送完一号空间目标或离轨碎片的瞬时转动惯量相等；
    (2)直线伸缩装置执行伸缩作业，将航天器系统的质心调整至待投送二号空间目标或离轨碎片的投送连杆(7)绕主连接轴(3)垂直旋转的投送平面内；当达到投送要求时，即投送二号空间目标或离轨碎片；
    (3)调整投送完二号空间目标或离轨碎片的投送连杆(7)上两个质量块(72)的位置，将该投送连杆(7)的质心调回至主连接轴(3)上，且该投送连杆(7)绕主连接轴(3)旋转的转动惯量与投送完二号空间目标或离轨碎片的瞬时转动惯量相等。
  9. 根据权利要求7所述的航天器章动抑制方法，其特征在于，所述步骤S5中消能卸载的具体步骤为采用正交强磁矩生成装置(5)产生的磁力矩反向作用于持续旋转的投送连杆(7)的转动惯量。
  10. 根据权利要求8所述的航天器章动抑制方法，其特征在于，所述步骤S5中消能卸载的具体步骤为采用正交强磁矩生成装置(5)产生的磁力矩反向作用于两个持续对向旋转的投送连杆(7)的剩余转动惯量。
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