45 results about "Magnetorquer" patented technology

The present invention provides a pure three-axis magnetic control preset performance attitude control method for an on-orbit spacecraft, comprising the following steps: Step 1: providing a spacecraft control torque calculation formula when a single magnetic torquer is used as an actuator; Step 2: providing a mathematical model of an attitude tracking system; Step 3: designing an attitude controller with preset performance to complete the attitude control with preset performance. The beneficial effects of the present invention are as follows: the pure three-axis magnetic control preset performance attitude control method for an on-orbit spacecraft of the present invention only uses a magnetic torquer as an actuator, and considering the requirements for static and dynamic performance required by the spacecraft in an actual working environment, a preset performance controller is designed to achieve the control of the dynamic steady-state performance of the attitude response of the spacecraft, thereby improving the safety and performance reliability of the magnetic control spacecraft system, and having broad application prospects in the fields of spacecraft docking and separation, on-orbit assembly, formation flying, rescue of faulty spacecraft, and de-orbiting of abandoned spacecraft.

The invention relates to a magnetic control method for a single-axis sun-oriented stable attitude of an optical microsatellite, and relates to the technical field of aerospace. The magnetic control method comprises the following steps: step 1, modeling optical microsatellite attitude dynamics adopting magnetic control; 2, calculating the posture of the single axis to the sun; 3, calculating an attitude control expected moment; and 4, calculating the expected magnetic moment of the magnetic torquer. According to the magnetic control method for the single-axis sun-oriented stable attitude of the optical microsatellite, the single-axis sun-oriented stable attitude control of the satellite can be realized under the condition of extremely low power consumption, and the energy stability of a non-task long-term mode and the reliable operation of the satellite can be ensured.

The invention discloses a flywheel unloading method and device based on PID control and a storage medium. The flywheel unloading method comprises the steps that the angular momentum deviation between the angular momentum of a flywheel deployed on a satellite and a reference angular momentum is determined; determining a PID control item for performing angular momentum unloading on the flywheel according to the angular momentum deviation; determining a control moment corresponding to the angular momentum deviation through a PID control mode based on a PID control item; according to the control torque, the unloading magnetic torque used for conducting angular momentum unloading on the flywheel is determined; according to the included angle between the magnetic moment and the unloading magnetic moment of the multiple magnetic torquers, weight information among the multiple magnetic torquers is determined, and the weight information is used for indicating the magnetic moment distribution proportion among the multiple magnetic torquers; according to the unloading magnetic moment and the magnetic moment distribution proportion, the magnetic moments of the multiple magnetic torquer are determined; and according to the determined magnetic moment, angular momentum unloading is carried out on the flywheel through a plurality of magnetic torquers.

The invention provides a high-precision KBR phase center on-orbit calibration planning and control method which is suitable for load calibration planning and control of a satellite with a magnetic torquer and a thruster as attitude control execution mechanisms. As a satellite is only provided with a magnetic torquer and a thruster as executing mechanisms, accurate attitude tracking control similar to momentum wheel control is difficult to simply realize. The invention designs a satellite load calibration maneuvering planning and control strategy based on a magnetic torquer and a thruster. The strategy mainly comprises five algorithm modules: an attitude maneuver planning target angle, angular velocity and angular acceleration real-time calculation module, an attitude control error calculation module, an attitude control calculation and maneuver planning torque thruster distribution calculation module based on magnetic control and a thruster, and a control signal synthesis and output module. The five modules are organically combined, and the satellite load calibration maneuvering planning and control problem based on the magnetic torquer and the thruster is systematically solved.

The present invention is applicable to the field of aerospace control technology, and provides a method for attitude magnetic tracking control in satellite orbit transfer, the method comprising: controlling the satellite attitude by three orthogonal magnetic torquers installed along the +X, +Y and +Z axis directions of the satellite; obtaining the expected angle during the satellite orbit transfer, and when the actual angle is close to the expected angle, compensating the torque generated by thrust with the torque generated by gravity; tracking control of the satellite attitude is performed by using a preset sliding mode control law based on an auxiliary compensator and taking the difference between the actual magnetic control torque and the ideal magnetic control torque as the input of the auxiliary compensator; the present invention uses the magnetic torquer as an attitude control actuator in the satellite orbit transfer, plans the expected angle during the satellite orbit transfer, and can compensate the torque generated by thrust with the torque generated by gravity when the actual angle is close to the target angle, thereby ensuring the stability of the satellite system and reducing the energy consumption of the magnetic torquer.

A satellite attitude fault-tolerant control method based on a magnetic torquer comprises the steps that firstly, in each control period, according to a control signal of a previous beat and an attitude measurement signal of a current beat, magnetic torquer and thruster fault detection based on control measurement consistency scoring is carried out, and when a thruster fault is detected or a thruster fault signal is received, the thruster fault detection is carried out; under the condition that the backup exists, the backup thruster is switched; when a thruster fault is detected or a thruster fault signal is received and no backup is available, switching an attitude control mode into a pure magnetic torquer control mode; and when a fault of a certain magnetic torquer is detected or a fault signal of the magnetic torquer injected by other methods or the ground is received, the fault magnetic torquer is cut off, a backup magnetic torquer is introduced, and a magnetic torquer group configuration is reconstructed for distribution calculation of the magnetic torquer.

A dipole field magnet comprising an inner volume shielded from the magnetic fields of the magnet due to the current flowing in the magnet. The dipole field magnet may comprise a plurality of ring-shaped superconductor coils configured to shield an inner volume between the ring-shaped superconductor coils from the magnetic field of the dipole field magnet. The dipole field magnet may form part of a fusion reactor and / or a magnetorquer and / or other dipole field magnet application. In some cases the dipole field magnet may comprise a plurality of ring-shaped superconductor coils configured to shield an inner volume between the coils from magnetic fields and / or the dipole field magnet may create a low field region in the inner volume.

The application discloses a kind of ultra-low orbit satellite high-precision attitude control system and method, the system includes: attitude measurement sensitive module, attitude controller and control execution mechanism;Wherein, attitude measurement sensitive module is used to measure the attitude information of current satellite;Attitude controller is used to combine the attitude information of current satellite, using preset control strategy, the control amount input of control execution mechanism is solved;Control execution mechanism is used to generate the moment of influence satellite attitude movement according to control amount input, to control satellite attitude by reaction flywheel control and jet control alternate use mode.This application is based on propulsion system, reaction flywheel and magnetic moment ware, realizes the full interval high-precision high-stability attitude control target for ultra-low orbit satellite.

The application discloses a kind of micro-nano satellite angular momentum autonomous management method, belongs to satellite attitude control technical field.The application is in view of the problem that existing satellite angular momentum management adopts open loop method, and autonomous correction capability is poor.The application comprises: satellite is identified through on-orbit flywheel rotating speed data environment torque parameter autonomously, and parameter that changes speed is faster is updated autonomously along with task time extension;Only through the attitude bias of relative sun orientation and sail rotation, angular momentum accumulation inhibition and angular momentum unloading two task objectives can be realized by manipulating environment torque, without active control torque intervention such as magnetic torque device, jet unloading;The whole process does not need artificial intervention, and angular momentum autonomous management is realized.The application can be autonomously carried out when satellite has no task, and reliability is good.

A method of controlling a satellite using only magnetic forces, and a control system for implementing the method. The method involves: using magnetic force measurements to assess a current attitude of the satellite at a current time and a current location; setting a desired attitude of the satellite at a future time at a future location; developing a set of waypoints that provide attitudes of the satellite at a plurality of locations between the current location and the future location; and actuating a plurality of magnetic torquers to cause torques that cause the differences between the attitudes of the satellite between the waypoints to be as small as possible; and achieving the desired attitude at the future location, the magnetic torquers being the only means of causing rotation of the satellite to achieve the desired attitude.

The invention relates to a satellite attitude safety capture control method based on magnetic measurement and magnetic control, which designs an angular velocity estimation algorithm depending on magnetometer measurement and a rate damping and sun orientation control algorithm only using a magnetic torquer as an execution mechanism under the condition that a gyroscope fails or air injection is unavailable. Satellite safety is guaranteed under the condition that the gyroscope breaks down or air injection is unavailable.

The invention relates to a high-precision magnetic torquer control circuit capable of realizing cooperative control with a magnetometer. The high-precision magnetic torquer control circuit comprises a cooperative control module, a magnetic torquer rapid turn-off control module, a magnetic torquer main driving circuit module and a magnetometer acquisition module. Working time slices of the magnetic torquer and the magnetometer are distributed in real time according to needs through the cooperative control module of the magnetic torquer and the magnetometer; rapid stabilization and zero returning of driving current of the magnetic torquer are realized through the rapid turn-off control module of the magnetic torquer; the output magnetic moment adjustment of the magnetic torquer is realized through the magnetic torquer main driving circuit module; and rapid real-time measurement of the spatial magnetic field intensity and the vector is realized through the magnetometer acquisition module. The cooperative control of the magnetic torquer and the magnetometer is realized, so that the idle time slice of the driving element and the detection element is shortened under the condition that the magnetic field is completely decoupled, and the detection efficiency and the control stability of the ultra-static satellite are greatly improved.

The utility model relates to spaceflight equipment maintenance tool technical field especially relates to magnetic moment ware coil quick replacement device. Including winding seat and the coil that winds outside winding seat, the protruding piece is integrally connected with both sides in winding seat one end, the fixed groove is equipped with inside both sides in winding seat other end, the fixed groove is equipped with with the limiting stopper of winding seat sliding connection and limiting stopper extends to winding seat outside, limiting stopper one side is connected with winding seat through a plurality of first spring and a plurality of first spring all are located in fixed groove, the fixed groove one side in winding seat is equipped with installation groove, the limiting fixed mechanism that is equipped with with limiting stopper cooperation in installation groove. The utility model provides a device that can significantly simplify operation process, can realize coil quick replacement to improve maintenance efficiency and reduce operation risk without completely disassembling.

The application discloses a kind of unloading method and device of asymmetric configuration satellite angular momentum, the method includes: by solar sensor, the solar vector under satellite body system is acquired, and the solar vector is converted into the solar vector under orbit system;The included angle β of the solar vector under the orbit system and orbit plane is calculated, and the rotation angle of solar wing rotation axis and the preset angle of biasing axis are set according to the included angle β of the solar vector under the orbit system and orbit plane;The angular momentum of flywheel system is calculated and is converted into the angular momentum under satellite body system;According to the modulus of satellite to be unloaded angular momentum and preset unloading threshold value are compared, according to the comparison result, it is judged whether magnetic moment device unloading needs to be started, and according to the comparison result, the control instruction of magnetic moment device is output.The application effectively reduces angular momentum accumulation and space moment accumulation;With the advantages of saving fuel, low power consumption requirement and high attitude stability, it provides a theoretical basis for angular momentum unloading of weak magnetic field planet explorer.

This invention discloses a method for analyzing the control capability of a satellite magnetic torquer, belonging to the field of satellite attitude and orbit control technology. The method first acquires the three-axis magnetic field strength data of the satellite's own system; then, it iterates through all specified torque directions. For each specified torque direction, it first calculates the average value of the maximum magnetic torque under different magnetic field strengths, and then calculates the corresponding magnetic torquer control capability; finally, based on the magnetic torquer control capability in each specified torque direction, it analyzes whether the magnetic torquer meets the control requirements. This invention can accurately obtain the control capability of the magnetic torquer, providing a decision-making basis for the selection and configuration of magnetic torquers.

The application provides a spacecraft pure magnetic control sun safety control method and system based on a magnetic moment device, which comprises the following steps: calculating a normal vector of a rotation shaft plane through a sun vector and a sun-pointing vector of the system, normalizing a magnetic field vector and constructing a unit orthogonal vector in the normal plane, solving an optimal rotation shaft vector in combination with a rotation inertia array, dynamically adjusting a control parameter according to a geometric included angle between the optimal rotation shaft and an inertial angular velocity of the system, superimposing a magnetic control sun pointing directional control amount and a damping control amount to obtain a total control amount. The system comprises vector calculation, optimal rotation shaft solving, control parameter adjustment and magnetic control amount output modules connected in sequence, and can be additionally provided with an execution driving module. The application fully considers the geometric relationship of multiple vectors, avoids attitude oscillation divergence, and can realize a sun pointing directional accuracy of ≤15° in less than one orbit period, thereby guaranteeing the charging of a solar cell array and providing a reliable pure magnetic control scheme for a spacecraft safety mode, and the engineering application value is remarkable.

The present invention discloses a method for using a magnetometer in a strong magnetic environment, which includes the following steps: Step 1, preprocessing of the magnetometer and validity judgment; Step 2, calculating the magnetic field intensity of the aircraft body relative to the inertial system according to the magnetic field model; Step 3, selecting the magnetic field reference; Step 4, determining the usage strategy according to the selection situation of the magnetic field reference. This method is based on reserving an interface for deducting the strong magnetic constant value from the magnetometer measurement value and the time-division and staggered control of the magnetometer and the magnetic torquer, which can avoid the problem that the magnetic field measurement value of the magnetometer is affected by the surrounding strong magnetic environment and its access to the closed-loop use, and can more reliably introduce the geomagnetic field value into the closed-loop control for use. The design of this method for using the magnetometer is simple and easy to be applied in engineering.

The invention relates to the technical field of spacecraft attitude control, in particular to a hollow-core magnetic torquer for a spacecraft, which comprises a coil body, the coil body comprises an inner-layer coil, a middle-layer coil and an outer-layer coil, a base is mounted at the bottom of the outer-layer coil, and screw rods are mounted at the positions, close to the two ends, of the top of the base through shaft seats. The inner-layer coil is connected with the lead screw at one end of the base, the outer-layer coil is connected with the lead screw at the other end of the base, and a motor driving the lead screws to rotate is mounted at the top of the base and corresponds to the lead screws. According to the invention, the coils are arranged into three groups, and two groups of coils are arranged to be capable of relatively axially moving, so that the magnetic field intensity can be indirectly changed by changing the magnetic field superposition amount of the coils on the premise that the input current is fixed, and therefore, the magnitude of the output torque can be changed, and compared with the existing pulse width modulation control mode, the output torque can be changed. An additional control mode can be supplemented and replaced, and a standby control scheme is provided for the magnetic torquer.

The application provides a cross-space satellite attitude control strategy autonomous switching method based on environment estimation, respectively calculates a main angular momentum increment, a coupling angular momentum increment and a reaction wheel group angular momentum increment, estimates a satellite environment disturbance torque by using the main angular momentum increment, the coupling angular momentum increment and the reaction wheel group angular momentum increment, and autonomously switches a control strategy according to the estimated value of the satellite environment disturbance torque; wherein, taking a continuous orbit time as a period, if a current orbit satellite environment disturbance torque three-axis angular momentum accumulation exceeds a current reaction wheel residual angular momentum capacity, and the satellite environment disturbance torque is greater than half of a magnetic torque controller control torque, the jet control strategy is autonomously switched to. The application can avoid early switching into the jet control, reduce fuel consumption, and on the other hand, avoid late switching out of the wheel control, so as to cause the wheel angular momentum saturation and satellite attitude divergence.

The invention relates to an in-orbit angular momentum adjusting method and device for a limited frame angle CMG. The method comprises the following steps: acquiring a spacecraft angular velocity, a CMG structure parameter and a control moment in a current control period; solving a magnitude mapping relation among the parameters according to a robust pseudo-inverse algorithm to obtain a frame angular velocity instruction; in the rapid maneuvering stage, under the constraint of the maximum frame angular velocity, an idle switch judgment result is obtained according to the analysis logic based on the frame angular deviation so as to correct a frame angular velocity instruction, and therefore the position and posture of the spacecraft are adjusted under the condition that the frame angular motion path is optimized; and if in the steady-state operation stage, a control magnetic moment instruction of the magnetic torquer is calculated according to the analysis logic based on the frame angular velocity deviation, and the frame angle is driven to reset by combining the control magnetic moment instruction and the frame angular velocity instruction. By the adoption of the method, it can be ensured that the spacecraft can keep high-precision attitude control in the rapid maneuvering stage and the stable operation stage.

The present invention discloses a method and system for controlling orbital maneuver divergence during the gravity satellite center of mass calibration process, relating to the field of spacecraft control technology. The method comprises: configuring the output of a magnetic torquer so that the output of the magnetic torquer is converted from an angular acceleration signal to an initial square wave signal or an initial sine wave signal; setting the phase of the square wave signal or the initial sine wave signal at the initial moment to 1 / 4 of a cycle to obtain a square wave signal or a sine wave signal; and performing center of mass calibration on the gravity satellite based on the method signal or the sine wave signal. The present invention solves the problem of orbital maneuver divergence during the gravity satellite center of mass calibration process and prevents the gravity satellite from losing control during the center of mass calibration process.

A three-axis integrated intelligent attitude control component suitable for a micro-satellite has a simple design and can independently design the system without having to redesign the system according to different satellites. It occupies a small space, is relatively light in weight, and has high system reliability. The component comprises: an X-axis reaction flywheel (11), a Y-axis reaction flywheel (12), a Z-axis reaction flywheel (13), an X-axis magnetic torque device (21), a Y-axis magnetic torque device (22), a Z-axis magnetic torque device (23), an attitude sensor (3), a mounting frame (4), an attitude control box (5), and a magnetic torque controller (6). The number of the X-axis reaction flywheel, the Y-axis reaction flywheel, and the Z-axis reaction flywheel is three, and the height is smaller than the diameter. The X-axis magnetic torque device and the Y-axis magnetic torque device are both long magnetic torque devices, and the Z-axis magnetic torque device comprises a plurality of short magnetic torque devices (231).

The application discloses a flywheel unloading method and device based on PID control and a storage medium. The method comprises the following steps: determining an angular momentum deviation between an angular momentum of a flywheel arranged on a satellite and a reference angular momentum; determining a PID control item for angular momentum unloading of the flywheel according to the angular momentum deviation; determining a control torque corresponding to the angular momentum deviation in a PID control mode based on the PID control item; determining an unloading magnetic moment for angular momentum unloading of the flywheel according to the control torque; determining weight information between a plurality of magnetic moment devices according to an included angle between a magnetic moment magnetic moment of the plurality of magnetic moment devices and the unloading magnetic moment, wherein the weight information is used for indicating a magnetic moment distribution ratio between the magnetic moment devices; determining the magnetic moment magnetic moment of the plurality of magnetic moment devices according to the unloading magnetic moment and the magnetic moment distribution ratio; and performing angular momentum unloading on the flywheel by the plurality of magnetic moment devices according to the determined magnetic moment magnetic moment.

The utility model relates to a magnetic moment measuring device of a magnetic torquer. The device comprises a magnetic moment test bench, and a two-dimensional moving support is arranged on the magnetic moment test bench. A plurality of magnetometers are arranged on the two-dimensional moving bracket at intervals, and each magnetometer is fixed on the two-dimensional moving bracket in a sliding manner; the magnetic moment test bench is also provided with a test rotary table, and a to-be-tested magnetic torquer is disposed on the test rotary table. The test rotary table is arranged on one side close to the magnetometers, and the magnetic induction intensity of the to-be-tested magnetorquer is measured through the magnetometers. The device provided by the utility model can be suitable for the measurement of the to-be-measured magnetorquer with different lengths, is simple in operation, is high in measurement precision, is high in accuracy, and is high in adaptability.

The present invention discloses a magnetic torque actuator drive and telemetry acquisition circuit for spacecraft, which is used to convert the received digital remote control signal into an analog voltage and output it to the drive control circuit of the drive power amplifier circuit. It is used to receive the analog voltage and output a corresponding drive current to the external drive power amplifier circuit according to the level of the voltage in proportion. It is used to collect the magnitude and direction of the drive current output by the drive power amplifier circuit to judge whether the response to the digital remote control signal is correct, namely the drive current telemetry acquisition circuit. The present invention uses a Darlington drive circuit structure composed of composite triodes to improve the drive ability and stability of the output current. By using the "virtual short" and "virtual open" characteristics of the input terminal of the operational amplifier, the control accuracy of the output current is improved. The structure of two zener diodes in parallel is adopted, so that the voltage at both ends of the output terminal can be controlled within a safe range. The real-time and accurate monitoring of the current can be realized by using the current sampling resistor to output the current telemetry through the triode.

1. Name of the designed product: magnetic torque device. 2. Use of the designed product: for satellite attitude control. 3. Design points of the designed product: in shape. 4. Picture or photo best indicating the design points: perspective view.

The invention provides a universal satellite attitude and orbit control test system self-checking device and an implementation method thereof. Comprising the steps of isolation measuring point resistance value detection, flywheel signal simulation, electromagnetic valve detection, magnetic torquer current signal simulation, sun sensor unipolar constant current source acquisition, magnetometer signal acquisition, watchdog reset and star sensor enable signal acquisition, serial port communication detection and electric star mode communication detection. The satellite-borne software communication detection and electromagnetic valve self-locking valve resistance value detection module realizes simulation of performance indexes required by testing of satellite attitude and orbit control satellite-borne products such as a flywheel, a sun sensor, a fixed star sensor, a propulsion electromagnetic valve self-locking valve, a magnetic torquer and a magnetometer. And the attitude and orbit control self-checking device is connected with the attitude and orbit control ground test system cabinet through a standardized connector and a matched cable.

The invention discloses a magnetic torquer magnetic current driving real-time compensation method, which comprises the following steps: firstly, carrying out voltage reverse acquisition on a magnetic current driving circuit, calibrating the precision of acquired resistance and voltage, then recording the magnitude of a magnetic current instruction and an actual magnetic current within a period of time, and carrying out linear fitting through least square to obtain a magnetic current driving real-time compensation result; and a magnetic current compensation coefficient is calculated, and real-time correction of the magnetic current driven by the magnetic current is realized. According to the method, the actual magnetic current driven by the magnetic torquer can be calculated and compensated on orbit in real time, high-precision control over the magnetic current is achieved, the precision of the output torque of the magnetic torquer is improved, and a better foundation is provided for satellite high-precision attitude control.

The invention discloses a satellite magnetic torquer control capability analysis method, and belongs to the technical field of satellite attitude and orbit control. The method comprises the following steps: firstly, acquiring triaxial magnetic field intensity data under a satellite body system; then, traversing all specified torque directions, for each specified torque direction, firstly calculating the average value of the maximum magnetic torques under different magnetic field intensities, and then calculating the corresponding magnetic torquer control capability; and finally, based on the control capability of the magnetic torquer in each specified torque direction, analyzing whether the magnetic torquer meets the control requirement or not. According to the method, the control capability of the magnetic torquer can be accurately obtained, and a decision basis is provided for configuration selection of the magnetic torquer.

A method for performing autonomous fault diagnosis on a magnetic torquer by using a correlation method belongs to the technical field of spacecraft attitude control, and comprises the following steps: (1) establishing a relationship between control voltage and current telemetry signals of the magnetic torquer; (2) designing constraint conditions of a magnetic torquer control and acquisition time sequence; and (3) extracting, counting and carrying out correlation fault diagnosis. The method disclosed by the invention is high in universality, and can be used for autonomously diagnosing faults of a control system provided with a magnetic torquer; through on-orbit statistical analysis of the correlation between the input and output of the magnetic torquer, the fault diagnosis problem of the magnetic torquer is solved, and the stability and safety of satellite operation are enhanced.