92 results about "Magnetorquer" patented technology

The invention provides a stable control system and method for a satellite spinning around the sun. A simulation sun sensor is used for measuring a sun vector under a satellite body system; a triaxial magnetometer is used for measuring a terrestrial magnetism vector under a satellite body; a controller is electrically connected with the simulation sun sensor and the triaxial magnetometer and used for obtaining the sun vector and the terrestrial magnetism vector, judging the current mode of the satellite and generating and sending a control command according to the current mode; and the magnetic torquer is used for outputting the corresponding magnetic torque according to the control command sent by the controller. By the adoption of the system and method, the satellite can stably spin around the sun through the high-reliability and low-power-consumption execution component magnetic torquer, and the satellite is passively stabilized and obtains energy. A guarantee is provided for the reliability and the safety of the satellite, and the defect that large-power-consumption and consumptive components such as a propeller and a flywheel are adopted in a satellite in the prior art is overcome.

The invention discloses an optical imaging minisatellite attitude control system and an in-orbit working mode switching method, belongs to the field of satellite attitude control and aims at solving the problem that existing optical imaging satellites cannot be simply and effectively switched in the orbit. The optical imaging minisatellite attitude control system comprises an attitude measurement sensor, an actuating mechanism and an attitude controller, wherein the attitude measurement sensor comprises a sun sensor, a star sensor and a gyroscope; the actuating mechanism comprises a counteraction flywheel and a magnetorquer. According to an in-orbit working mode switching method of the optical imaging minisatellite attitude control system, the optical imaging minisatellite attitude control system has six working modes, namely a rate damping mode, a sun capture mode, a sun-orientated three-axis stabilization mode, an earth-orientated three-axis stabilization mode, a data transmission mode and a safety mode. The optical imaging minisatellite attitude control system and the in-orbit working mode switching method are suitable for all the optical imaging minisatellites.

The invention discloses a novel space station extravehicular inspection maintenance unit (EIMU), and belongs to the design field of aerospace crafts. The novel space station EIMU comprises intravehicular monitoring equipment, an antenna and an EIMU device, wherein the intravehicular monitoring equipment and the EIMU device are in communication through the antenna. The intravehicular monitoring equipment comprises a video display terminal, an inspection device control platform and a manual remote control device. The EIMU device comprises solar cell panels, a micro-mechanical arm, sun sensors, nitrogen propellers, a binocular stereo vision camera, a three-axis magnetorquer, a momentum wheel, a satellite-borne computer board and a communication board. The novel space station EIMU can provide extravehicular inspection and maintenance service for large spacecrafts such as space stations and ensure that the spacecrafts can achieve in-orbit operation safely and fulfill tasks smoothly.

The invention relates to a high-energy-efficiency hollow magnetic torquer design method and manufacturing process. The basis physical principle than magnetic force acts on an electrified coil in a magnetic field, the magnetic torque of a hollow magnetic torquer is optimized under the constraint conditions of size, quality and power consumption of the hollow magnetic torquer, and the manufactured magnetic torquer is lightweight, low in energy consumption, small in residue magnetism and large in output magnetic torque. The requirements of the posture control system of a near-earth orbit modern small satellite for magnetic torquers can be satisfied. Efficient magnetic torquers under different constraint conditions can be manufactured by the method. Traditional magnetic torquers using magnetic bars are large in magnetic torque, but high in magnetic core material requirement and large in residual magnetism. The method uses a hollow coil, the defect of large residual magnetism is overcome. Due to the fact no magnetic cores are used, the mass of the magnetic torquer is reduced greatly.

The invention relates to a flexible support based torque calibration and measurement device. Left and right scale pans used for hanging of weights are symmetrically mounted on two sides of a standard torque beam through elastic pieces. A rotation center 'O' of a flexible support coincides with a central axis 'Y' of the standard torque beam. A laser displacement sensor mounted on a base is arranged below the lower end face of the standard torque beam. Left and right electromagnetic torquers are symmetrically mounted on two sides of the lower end face of the standard torque beam. On the condition of a torque load, the laser displacement sensor transmits measured electric signals of displacement variation, of a laser sensor probe and a standard torque lower end face, caused by the torque load to the left electromagnetic torquer or the right electromagnetic torquer after the electric signals are processed by a measuring circuit so as to control the left electromagnetic torquer or the right electromagnetic torquer to generate electromagnetic force to level the standard torque beam. The flexible support based torque calibration and measurement device has the advantage that measurement precision is improved to the utmost by comprehensive measures such as an electromagnetic force equilibrium system, the high-precision laser displacement sensor and a rapid signal tracking and measuring system.

The invention discloses a cable-free high-density cubesat and an assembly method thereof. The cable-free high-density cubesat comprises a main structural frame 1, a solar cell array 2, an antenna plate 3, an on-satellite integrated electronic system 4, a momentum wheel 5, a gyroscope 6, a magnetorquer 7, a remote sensing camera 8, a battery pack 9, a camera fastening frame 10 and a lead screw 11.The main structural frame 1 includes an upper end cover 101 and a satellite shell 102, and the momentum wheel 5, the gyroscope 6, the magnetorquer 7 and the camera fastening frame 10 are fixed into the satellite shell 102; and the remote sensing camera 8 is fixed into the camera fastening frame 10, the on-satellite integrated electronic system 4 is a multi-layer circuit board formed by folding a whole circuit board, the multi-layer circuit board is placed in the satellite shell 102, and the solar cell array 2 is fixed to the side wall of the main structural frame 1. The cable-free high-densitycubesat realizes the on-satellite integrated electronic integration and the non-cable between systems, realizes the high-density integration of a cubesat, and improves the utilization ratio of the internal space of the cubesat.

The invention discloses a method for controlling the procession of a bias momentum micro-satellite, which comprises the following steps: performing initialization: the Y-axis output of a magnetometer at the current moment is acquired and recorded, and model geomagnetic field of the satellite in an orbital coordinate system is calculated and recorded referring to a geomagnetic table and according to the current time as well as the orbit parameter; reading out the Y-axis output of the magnetometer recorded at the previous cycle; obtaining A (see details in the specification) by Y-axis outputting the magnetometer at the current moment and subtracting the magnetometer Y-axis output recorded at the previous cycle; obtaining B (see details in the specification) by subtracting the Y-axis model magnetic field value at the previous cycle from that in the orbital coordinate system at the current moment; according to a control law C (see details in the specification), calculating out the output magnetic torque of a Y-axis magnetic torquer, and sending control command to the magnetic torquer; recording the Y-axis output of the magnetometer at the current moment; recording the Y-axis model magnetic field value in the orbital coordinate system at the current moment; and controlling to repeat steps 2) to 9) taking the cycle T as uniform intervals. The invention has the advantages that the measurement to attitude angle is omitted, the operation amount is less, and the dependence on hardware is reduced.

The invention discloses a single-piece three-axis magnetorquer, and belongs to the technical field of satellite attitude control. The single-piece three-axis magnetorquer comprises three electrification loops in the orthogonal direction and corresponding driving control circuits, wherein the two electrification loops in the orthogonal direction are respectively magnetic bars generating the magnetic moment under the effect of driving current, the third electrification loop in the orthogonal direction is a hollow coil generating the magnetic moment under the effect of driving current, the hollow coil and the two magnetic rods are integrated on a coil circuit board, three driving circuits are integrated on a driving circuit board, the driving circuit board is connected with a coil circuit board through a standard row pin, the two magnetic bars are vertically integrated on the circuit board for forming a Descarte rectangular coordinate system, n layers of hollow coils are wound on the coil circuit board, and a hollow coil array with the wire diameter being d is formed. The single-piece three-axis magnetorquer has the advantages that the coil circuit and the driving control circuit are integrated on the circuit board, in addition, two circuit boards are conducted through a pointer, the magnetorquer is realized by the single-piece structure, and the structure is simplified.

The invention discloses angular momentum control method of a spacecraft in an inertial system. The method is characterized in that a three-axis stabile control is performed in the inertial system, and the feedback design is performed based on the angular momentum and three-axis posture angle information of a performing mechanism to obtain a feedback controller. According to the method, that the gravity gradient torque is to be more than the pneumatic torque is removed; the performing mechanism is unsaturated while the long-term stable flying is ensured; the method is applied to the spacecraft posture control in the circumstance that a magnetorquer cannot be assembled or the magnetorquer suffers from fault and unloading cannot be or is not proposed to be performed by the air injection mode.

The invention relates to a high-precision radiation-resistant microsatellite three-axis magnetorquer which comprises two magnetic rods provided with metal cores, a hollow magnetorquer body, a processor and magnetometers. The two magnetic rods provided with the metal cores form an X-axis magnetic rod and a Y-axis magnetic rod respectively and are wound by coils, the hollow magnetorquer body forms a Z-axis hollow coil, the X-axis magnetic rod, the Y-axis magnetic rod, the Z-axis hollow coil, the processor and the magnetometers are integrated into a PCB and form modularization, and three execution mechanisms are independently driven by PWM signals. The invention further relates to a work method of the high-precision radiation-resistant microsatellite three-axis magnetorquer. The high-precision radiation-resistant microsatellite three-axis magnetorquer and the work method thereof have the advantages that reliability of the magnetorquer is improved, and control precision of the magnetorquer is improved; the functions of current monitoring and temperature measuring are achieved, the magnetic rods and a fastening piece are integrally designed, the magnetorquer is convenient to assemble, and strength is increased.

The invention provides a satellite integrated controller. The controller comprises a computer module and an interface module; the computer module integrates an attitude and orbit control computer, a house-keeping computer and a calculation and data processing unit of a lower computer of a power supply controller, and corresponding computer functions are achieved through the computer module uniformly; the interface module is used for providing the interface function needed by communications of the computer module and external devices; the external device comprises one or more of a sun sensor, astar sensor, a magnetometer, a gyroscope, a flywheel, a magnetorquer and the power supply controller. Accordingly, integration design is conducted on an existing satellite platform resource, design of a satellite platform is simplified, and the satellite platform resource is saved.

The invention provides a satellite attitude control magnetic torque; the satellite attitude control magnetic torque comprises a control computer, a magnetic wiping control circuit and a permanent magnet body; the control computer is used for transmitting a magnetic torque write-in signal and a magnetic torque wiping signal; the magnetic torque wiping signal writes the permanent magnet body in a rated magnetic torque when the magnetic torque write-in signal is received, and wiping off the magnetism of the permanent magnet body when the magnetic torque wiping signal is received. The satellite attitude control magnetic torque greatly reduces time required to complete gesture control while saves more energy sources, and is more reliable. The invention further provides a satellite attitude control system and a satellite.

The invention discloses an electromagnetically-driven grating accelerometer and a closed-loop detection method. The method comprises the steps of achieving the photoelectric conversion of a diffraction interference signal outputted by an accelerometer sensitive structure through a photoelectric detector, enabling the diffraction interference signal to enter a digital detection circuit after beingamplified and filtered through a pre-amplification filter circuit, and carrying out the analog-to-digital conversion through an A / D converter, performing phase-sensitive demodulation on a demodulationsignal generated by an FPGA module in a phase-sensitive detection module, and filtering noise and background stray light interference through a digital low-pass filter to obtain a purified closed-loop error signal, sending the purified closed-loop error signal to a controller for PID control adjustment to obtain a feedback signal, on one hand, enabling the feedback signal to serve as output, on the other hand, enabling the feedback signal and a modulation signal generated by the FPGA module to be converted through a D / A converter, and after signal amplification is conducted through a linear amplification circuit, loading the feedback signal on a magnetic torquer of a sensitive structure; phase modulation of a cantilever beam is carried out by changing the adsorption effect of a mass blockand a magnet, and servo balance is achieved.

The invention discloses a measuring coil of a spacecraft magnetorquer, wherein the measuring coil comprises a coil framework, a coil winding and a wire-connecting column; the coil framework comprises two parallel opposite high-low temperature resistant material plates which are symmetrical in shape; the shape of the upper part of each plate is more than half roundness; the lower part of the roundness forms a long rectangular shape through a transition circular arc; a cylinder is connected and arranged between the two plates; a through hole is formed on the two plates and the cylinder along the common circle centre axis in a penetrating manner; the inside diameter of the through hole is slightly more than the outside diameter of a magnetorquer to be measured; the coil winding is wound along the outside diameter of the cylinder between the two plates of the coil framework; the wire-connecting column is fixed at two rectangular sides at the lower part of the two plates of the coil framework; and two ends of the coil winding are connected to the wire-connecting column. The invention also discloses a measuring method by utilizing the coil. Compared with the prior art, the method disclosed by the invention has the advantages of being rapid in measurement speed, simple in measurement step and easy to operate and perform automatic test; in addition, automatic testing equipment is very easy to develop according to the method.

The invention provides a high-density modular magnetic torquer applied to a micro-nano satellite and a design method, and relates to the field of control of magnetic torquers. The problems that the existing on-satellite magnetic torquer is large in size, complex in cables and excessive in circuit boards involved, and thus other on-satellite devices are affected are solved. The magnetic torquer comprises an X-axis magnetic bar, a Y-axis magnetic bar, a Z-axis hollow magnetic torquer and a PCB. The X-axis magnetic bar and the Y-axis magnetic bar are perpendicularly mounted on the PCB. The hollow magnetic torquer is composed of a hollow square coil and integrated on the PCB. The length and diameter of the X-axis magnetic bar and the length and diameter of the Y-axis magnetic bar are obtained according to magnetic torque indexes of the X-axis magnetic bar and the Y-axis magnetic bar; the size of the PCB is determined; the length and width, on the PCB, of the trace of the Z-axis hollow magnetic torquer are further obtained; finally the Z-axis hollow magnetic torquer is arranged on the PCB, and thus the design of the modular magnetic torquer is realized. The method has the advantages of high operability, simple external interfaces, and simplification of on-satellite cables, and the launching cost of the satellite is greatly reduced.

The invention provides an integrated micro-nano satellite electronic system. The integrated micro-nano satellite electronic system comprises a satellite core electronic unit, a solar wing driving mechanism, a three-axis gyroscope, a three-axis magnetometer, a simulated solar goniometer, a three-axis integrated micro flywheel, a magnetic torquer, a thermistor, a heater, an initiating explosive device, a measurement and control antenna, a GNSS antenna and a data transmission antenna. The satellite core electronic unit is provided with any one or more of the following interfaces: an analog quantity acquisition interface, a motor drive interface, an initiating explosive device driving interface, a heater drive interface, a magnetotorquer drive interface, a CAN bus interface, a high-speed datatransmission interface, and a radio frequency interface. A traditional satellite subsystem type integration mode is replaced, a satellite core electronic unit serves as an information and energy center, components with other functions are matched, satellite composition is simplified, and miniaturization, light weight and high integration design of the micro-nano satellite is facilitated.

The invention provides a modular attitude control unit for a satellite. The modular attitude control unit comprises a shell, a momentum wheel and a magnetic torquer, wherein the surface of the shell comprises at least one standard interface; the modular attitude control unit is connected to other functional modules of the satellite through the standard interface; the momentum wheel and the magnetic torquer are arranged in the shell; the momentum wheel is used for controlling the attitude of the satellite to maintain the angular momentum of the system; and the magnetic torquer is used for performing attitude control or momentum wheel unloading management on the satellite.

The satellite three-axis attitude control method combined with the magnetic torque device when the flywheel is underactuated, the steps are: (1) Determine the magnetic control matrix Γ(b) according to the geomagnetic field intensity vector B measured by the magnetometer, and according to the actual flywheel availability , determine the flywheel control matrix Kw; (2) calculate the magnetic current command Izk=Kic*Mzk for three-axis attitude control; (3) calculate the magnetic current command Ixz=Kic*Mxz for three-axis magnetic unloading; (4) calculate the magnetic torque device Command Ick=Izk+Ixz; while the magnetic torquer performs attitude control according to the magnetic torquer command, the available reaction flywheel continues to perform attitude control based on the command torque Tc. The method of the present invention realizes the high-precision control of the satellite's three-axis attitude in combination with the three-axis magnetic torque device when there are only two or even one flywheel available in the system, and is suitable for long-term ground-oriented control or long-term sun-oriented control, and also It is suitable for the maneuvering control of satellites at any attitude and large angle.

The invention relates to a multi-objective design method for a micro-star magnetic torquer. The method is based on a micro-star magnetic torquer multi-objective design system. The system comprises a parameter input unit (11), a magnetic core and winding database (12), a design unit (13), a screening unit (14) and a file output unit (15). The design unit (13) performs analysis and calculation according to related design information of the parameter input unit (11), the magnetic core and the magnetic torquer of the winding database (12) to obtain all feasible design schemes; the screening unit (14) screens an optimized design scheme from all feasible design schemes according to a given screening condition, and extracts an optimal design scheme in a screening condition range at the same time.The obtained result is output and displayed through a file output unit (15).

The invention provides a magnetic torquer for a microsatellite and an assembly method thereof, and belongs to the field of lightweight design technologies and assembly processes of the magnetic torquer. The magnetic torquer comprises a left mounting rack, a coil, a magnetic core, a right mounting rack, a left insulating baffle plate and a right insulating baffle plate; the two ends of the magnetic core are connected with the left installation frame and the right installation frame respectively, the coil is wound on the magnetic core, a left insulation blocking piece is arranged at the connecting position of the magnetic core and the left installation frame, and a right insulation blocking piece is arranged at the connecting position of the magnetic core and the right installation frame; the magnetic core is made of permalloy meeting the national standard 1J85, and the coil is wound by a single-layer polyimide enameled wire in a multi-layer and multi-turn mode. The magnetic torquer is designed in a light weight mode, a magnetic core material with low quality and low residual magnetism in national standard products and a single-layer enameled wire with the optimal scheme wire diameter are selected, an optimal winding scheme is adopted through a multi-target recognition method, the design structure is reduced, the weight of the magnetic torquer is reduced to the maximum extent, the optimal design is achieved, and therefore the satellite launching cost is reduced.

The invention provides a throwing cable type space flexible electromagnetic docking mechanism which is divided into three parts, a driving mechanism and a control system are mounted on a tracking spacecraft, and a driven mechanism is mounted on a target spacecraft. A throwing cable mechanism, an electromagnetic coil and a mechanical claw locking mechanism are mounted in the driving mechanism. The throwing cable mechanism throws out a guide ball so that the guide ball can be captured by the driven mechanism and a throwing cable is recovered to complete butt joint. Electromagnetic force generated by the electromagnetic coil serves as power for launching a throwing cable and buffering power before contact with spacecrafts. A mechanical claw locking mechanism is used for limiting the guide ball in the non-butt-joint period and used for locking the two spacecrafts after butt joint is completed. The control system comprises a laser distance measuring sensor, an attitude sensor, a magnetic torquer and a flywheel. The distance measuring sensor provides relative position information between the two spacecrafts so as to determine whether an active spacecraft and a passive spacecraft meet the butt joint distance requirement or not. The attitude sensor is used for judging the relative attitude of the two spacecrafts and ensuring that the guide ball can be captured. The driven mechanism is provided with the electromagnetic coil for adsorbing the guide ball.

The invention discloses a method for manufacturing a hollow magnetic torquer. The method comprises the following steps that a, the four corners of a square fixing plate are each detachably provided with a winding groove; b, a wire is wound around the four winding grooves to form hollow square coils, and the wire is coated with epoxy resin to be cured while being wound; c, after wire winding is completed and the wire is cured, the fixing plate and the four winding grooves are separated, and the four winding grooves and the cured coil wound on the winding grooves are obtained, that is to say, the hollow magnetic torquer is manufactured. The method for manufacturing the hollow magnetic torquer has the advantage of solving the problems that when a traditional method is used for manufacturing the hollow magnetic torquer, the process is complex, and the stability is poor.

The invention provides an area coverage control method for an offset momentum satellite, which comprises the following steps that a first momentum wheel, a second momentum wheel and a third momentum wheel are adopted to execute attitude control in an offset momentum control mode, the nominal angular momentum directions of the first momentum wheel and the second momentum wheel are in a YOZ plane of a satellite body system and are distributed on two sides of a Y axis, the included angle with the Y axis is 45 degrees; the nominal angular momentum direction of the third momentum wheel is in the Z axis and is a backup of the two V-shaped momentum wheels; and the three-axis axial magnetic torquer is used for unloading the momentum wheel.

The invention discloses a multi-redundancy satellite intelligent attitude control assembly. The assembly comprises a box body (1), reaction flywheel sets (2), a magnetic torquer set (3), a magnetic torquer coil (4), a control panel (5) and gyroscopes (6), and the box body is provided with a magnetic torquer coil mounting groove (11) and a control panel mounting groove (12); the reaction flywheel sets include a first reaction flywheel set, a second reaction flywheel set, a third reaction flywheel set and a fourth reaction flywheel set which are arranged at the four corners of the box body respectively, each reaction flywheel set comprises an X-axis reaction flywheel (21), a Y-axis reaction flywheel (22) and a Z-axis reaction flywheel (23), and the X-axis reaction flywheel and the Y-axis reaction flywheel are installed on the side wall of the box body; the magnetic torquer set is used for performing magnetic unloading on the reaction flywheels and comprises an X-axis magnetic torquer and a Y-axis magnetic torquer which are mounted on a bottom plate of the box body and are symmetrically distributed; the magnetic torquer coil is installed in the magnetic torquer coil mounting groove of the box body and is in the Z-axis direction; the control panel is arranged in the center of the box body and is mounted in the control panel mounting groove; and the three gyroscopes are all installed on the bottom plate in the box body. The invention also provides a working method of the assembly.

A space vehicle includes one or more magnetorquers operable to change an attitude of the space vehicle in an external magnetic field, each magnetorquer comprising a coil, and a switching circuit for short-circuiting the coil of at least one of the magnetorquers so that a closed electric circuit comprising said coil is formed, for damping tumbling motion of the space vehicle in the external magnetic field. The switching circuit is configured to short-circuit the coil of the at least one magnetorquer upon occurrence of a condition indicative of end-of-life or failure of the space vehicle. The application further relates to a corresponding method of operating a space vehicle.