90 results about "Magnetorquer" patented technology

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 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, comprising: a computer module and an interface module; the computer module integrates the calculation and data processing unit of the lower computer of the attitude and orbit control computer, the star computer, and the power supply and distribution controller The corresponding computer functions are uniformly realized by the computer module; the interface module is used to provide the interface function required for the computer module to communicate with external devices; the external devices include: sun sensors, star sensors, magnetic One or more of strong gauges, gyroscopes, flywheels, magnetic torque devices and power supply and distribution controllers. The present invention carries out an integrated design on the existing satellite platform resources, simplifies the design of the satellite platform, and saves the resources of the satellite platform.

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 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 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 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 provides a device and a method for realizing sun capture by zero momentums of micro-satellites. The device comprises an analog sun sensor, a magnetometer and a magnetic torquer, whereinthe analog sun sensor and the magnetometer are used for together determining the attitude of a satellite; and the magnetic torquer is used for providing a control moment in a rate damping process andused for enabling a normal axis of a solar array of the satellite to point to the vector direction of the sun as well as controlling an included angle between a normal line of the solar array of the satellite and the vector direction of the sun to be less than a preset threshold, thereby realizing sun capture by a zero momentum magnetic control mode of the satellite. The device and the method provided by the invention simplify the working mode of an attitude and orbit control system and improve the working reliability when the satellite enters an orbit in the early stage.

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.

The invention provides a magnetic torquer signal processing method and system of satellite attitude and orbit control comprehensive test equipment. The method comprises the following steps: a magnetic moment signal resolving step: converting collected magnetic torquer driving current into real-time magnetic moment; a local geomagnetic field determination step: determining the local geomagnetic field intensity according to the real-time position; a magnetic control moment resolving step: resolving a magnetic control moment according to the local geomagnetic field intensity and the real-time magnetic moment information; and a step of adding the magnetic control torque to the dynamic model: establishing an attitude dynamic model, and superposing the magnetic control torque to the attitude dynamic model. The method is suitable for signal processing requirements of magnetic torquer for different satellite attitude and orbit control subsystems, so that the universality of signal processing of the magnetic torquer of satellite attitude and orbit control comprehensive test equipment is improved.