76 results about "Satellite attitude control" patented technology

The invention discloses an energy scheduling method, system and device for a low-earth-orbit satellite and a medium. The method comprises the steps that the energy storage level is calculated according to the current generation power of the low-earth-orbit satellite, the current remaining electric quantity of an energy storage unit and the health degree of a battery; generating a future illumination time sequence according to the orbit parameters of the low-orbit satellite and the satellite attitude control data; according to the current operation state data of each device loaded on the low earth orbit satellite and a preset task list, generating corresponding task prediction energy consumption data arranged in a descending order according to the importance level through a historical energy consumption model; according to the energy storage level, the future illumination time sequence and the task prediction energy consumption data, calculating the dominant power supply amount, matching the energy consumption demand in the task prediction energy consumption data, and carrying out the preferential power supply of the equipment of the task item with the high importance level. And carrying out power supply adjustment on the other equipment by adopting at least one strategy of delaying to the illumination period, reducing the operation power and shortening the operation duration so as to realize efficient scheduling on the energy of the low-orbit satellite.

The invention relates to the technical field of satellite attitude control, in particular to a satellite-borne laser semi-autonomous pointing control method, electronic equipment and a storage medium. The method is completed by cooperation of a ground system and an on-satellite system, the ground system calculates a top target passing moment T * in advance, generates a unified laser pointing control instruction and injects the instruction into a satellite; and the satellite-borne system receives and analyzes the instruction, obtains parameters such as T * and beam selection, and then autonomously controls a specified load for work preparation. And the satellite-borne system autonomously calculates the optimal moment T1 and the target attitude when passing through the calibration field through high-precision orbit extrapolation based on the latest navigation data, finally autonomously executes attitude maneuver and laser pointing tasks, and autonomously returns after the tasks are completed. Instructions in a uniform format are generated and injected from the ground, ground operation is greatly simplified, efficient cooperation of ground decision triggering and on-satellite autonomous execution is achieved, pointing errors caused by orbit extrapolation are greatly reduced, the pointing precision and the task success rate are improved, and universality is good.

The invention relates to the technical field of satellite attitude control, in particular to a satellite-borne laser full-autonomous pointing control method, electronic equipment and a storage medium. The method comprises the following steps: firstly, presetting the position of a ground fixed calibration field, a laser light emitting phase and pointing information on a satellite; the on-satellite system autonomously carries out long-term orbit extrapolation, judges whether a satellite is about to pass through a calibration field, calculates to obtain a rough target moment T * and a day and night state thereof, and autonomously notifies a ground user. And the satellite autonomously controls work preparation of laser and camera loads according to T *. And at the moment before T *, the satellite-borne system autonomously solves the optimal moment T1 and the target attitude of passing through the calibration field through iterative extrapolation calculation and preferential selection based on high-precision navigation data. And finally, the satellite autonomously maneuvers to a target attitude, laser is controlled to accurately point to the calibration field at the T1 moment, and the satellite autonomously returns after the task is completed. According to the method, on-satellite autonomy from task planning to execution of the whole process is realized, and the autonomy, the working efficiency and the pointing precision of the satellite are greatly improved.

The invention discloses a parameter self-adaptive control method for a low-orbit satellite with a solar panel driving device. Belongs to the technical field of aerospace, and particularly relates to the technical field of low-orbit satellite parameter control. The satellite attitude control method solves the technical problems of overcoming the influence of the rotational inertia of the satellite and compensating the influence of the SADA vibration disturbance torque and the atmospheric resistance torque on the satellite attitude control precision. The method comprises the following steps: constructing a satellite attitude kinematics and dynamics equation in a quaternion form, and constructing a satellite attitude kinematics and dynamics equation for designing an attitude controller on the basis of the satellite attitude kinematics and dynamics equation; defining an error attitude kinematics and dynamics equation of the satellite through the attitude control error and the angle control error; carrying out segmentation conversion on a to-be-estimated parameter and a known state variable, and expressing the negative influence of a rotational inertia error on the attitude by matrix operation; designing the control torque of the attitude controller by adopting a mode of combining nonlinear feedback control with self-adaptive control; and updating the self-adaptive estimated value in the control parameters of the attitude controller by adopting a parameter self-adaptive law.

The invention discloses an optical satellite drift angle accurate calculation and attitude correction method and system, and relates to the technical field of satellite attitude control, and the method comprises the steps: calculating a prime unitary circle curvature radius according to a WGS84 ellipsoid parameter, and converting the geodetic coordinate of a ground target point into an ECEF rectangular coordinate; calculating Julian days based on UTC time, determining an age difference angle and a nutation angle, and converting an ECEF coordinate into a J2000 coordinate; calculating the projection of the earth rotation angular velocity in a J2000 coordinate system, and determining a relative velocity vector; determining an orthogonalized body coordinate system in combination with a unit vector from the target satellite to the ground target point; and projecting the relative velocity vector to a body coordinate system, calculating a drift angle and a double-rotation-sequence Euler angle, and correcting a real-time execution attitude by rotating a matrix around a Z axis. According to the method, high-precision image motion compensation can be realized, the imaging quality of a high-resolution optical satellite in a complex task scene is improved, and a technical foundation is laid for wide application of sub-meter remote sensing images.

The invention discloses a satellite attitude control method and device for propulsion system faults, and belongs to the field of satellite attitude control. The method comprises the following steps: initializing a software available state of a satellite control propulsion system according to satellite control computer software; initializing a hardware available state of the satellite control propulsion system according to a hardware interface of the system; and determining whether target data is successfully recovered when the satellite control computer is reset or cut off, if so, performing satellite attitude control according to the available state of the software, and otherwise, performing satellite attitude control according to the available state of the hardware. The method can effectively improve the safety and stability of satellite operation when the propulsion system breaks down.

The application relates to a high-dynamic on-line interference estimation method for a ring-scan load on a satellite. The method designs an on-line high-dynamic interference prediction method based on an observer aiming at the periodic characteristics of the interference, and realizes high-precision estimation of fast-changing interference. The high-precision estimated value of the interference torque is used as a feedforward item of the satellite attitude control torque, so that the load interference is effectively inhibited, and the pointing accuracy and stability of the satellite are ensured.

The present application relates to a kind of satellite load observation attitude control and data transmission, orbit control autonomous intelligent planning system and method, the system includes mission guide module, attitude control module, orbit control module and data transmission module;The attitude control module includes several thrusters for satellite attitude control, the orbit control module is used for satellite autonomous orbit control calculation, and carries out satellite orbit control, the data transmission module is used for satellite autonomous data transmission calculation, and satellite data is transmitted to ground;The mission guide module is in system control core, executes load observation task implementation strategy, according to priority reasonably arrange satellite's working time sequence, so that the satellite is intelligently switched between attitude control and data transmission to ground, orbit control, controls satellite autonomous intelligent operation in orbit.Autonomous intelligent planning, autonomous data transmission calculation, autonomous orbit control calculation are beneficial.

The invention relates to the technical field of satellite attitude and orbit control, in particular to a non-inductive coordination control method for quick zero passage of a momentum wheel under a redundant momentum wheel train. The method comprises the following steps: firstly, interpreting the change trend of the rotating speed of the momentum wheel according to the change condition of the rotating speed of the momentum wheel; secondly, judging the zero-crossing state of the momentum wheel according to the rotating speed change trend and the actually measured rotating speed of the momentum wheel; and finally, based on the characteristic of zero motion of the momentum wheel set, applying zero-crossing compensation torque to the momentum wheel in a zero-crossing state to enable the momentum wheel to quickly pass through a rotating speed zero-crossing interval, and applying reverse compensation torque to other momentum wheels to enable the resultant torque of the compensation torque of the momentum wheel set to be zero. The rotating speed of the momentum wheel quickly passes through a zero-crossing interval through a method of increasing zero-crossing compensation torque, and the influence of zero-crossing of the rotating speed of the momentum wheel on the satellite attitude control precision is reduced. The method can be popularized and applied to all satellites with zero-crossing momentum wheel rotating speed and high attitude control precision requirements.

The application belongs to the technical field of satellite attitude control, and relates to a large flexible spacecraft distributed fault-tolerant attitude and vibration control method, a large flexible spacecraft rigid-flexible coupling dynamics model with distributed actuators is established by using the Lagrange method, a hierarchical Stackelberg game control framework is established by taking an attitude control subsystem as a leader layer and a vibration suppression subsystem and a fault-tolerant control subsystem as follower layers, performance functions of the three subsystems are defined, and through game of spacecraft distributed fault-tolerant attitude and vibration control, strategy cooperation and dynamic balance between at least two control targets are realized; the application realizes three-axis attitude stabilization and low-frequency vibration suppression of a target spacecraft, and shows stronger fault-tolerant capability under node failure.

The application discloses a mode switching method applied to a variable rotating speed control moment gyro, which clearly defines multiple working modes of the control moment gyro, switching logic between behaviors of the control moment gyro in the modes and the modes, can realize non-control, moment control and backup reaction wheel functions of the control moment gyro, and can make the control moment gyro effectively applied in satellite attitude control.

The application discloses a satellite attitude control method and device based on mode conversion and a satellite, relates to the field of satellite control, and is characterized in that: a fine control mode, a coarse control mode and a sun capture mode used for controlling the attitude of the satellite and having different attitude determination accuracies are set in advance, the fine control mode is determined as a current target attitude control mode when a first remote control signal is received, the stop control mode is determined to be entered when it is determined that the satellite failure this time is a system out-of-control failure, the attitude of the satellite is regulated according to a second remote control signal when the second remote control signal sent by a ground remote control module is received, and the target attitude control mode is determined again according to the type of the failure carried component when it is determined that the satellite failure this time is a failure of the carried component used for attitude control, so that the satellite is switched to a new target attitude control mode. According to the scheme, the satellite can be switched between different modes according to different states of the satellite, reliable control of the attitude of the satellite is realized, the scheme has high flexibility, strong self-adaptability and is beneficial to practical application.

The utility model relates to a satellite three -axis integration reaction flywheel module relates to satellite attitude control technical field, including base module, non -magnetic motor, harmonic reducer, wheel disc, regular quadrangular prism and reaction flywheel, the middle part of the top of base module is equipped with round recess, the wheel disc is rotatably connected in the round recess, is equipped with the tooth on the circumference of wheel disc, the side of base module is connected with non -magnetic motor, the output shaft of non -magnetic motor is connected through harmonic reducer and the tooth of wheel disc, is used for driving wheel disc rotation, the top of wheel disc is connected with regular quadrangular prism, all are equipped with quick detachable interface on the four sides of regular quadrangular prism, the reaction flywheel is connected on the side of regular quadrangular prism through quick detachable interface. Advantages lie in: compared with traditional three motor systems more energy -conserving, reduce control coupling, be convenient for on -orbit maintenance or replacement, improve the heat dissipation efficiency, prolong the flywheel life.

The invention relates to the technical field of satellite attitude control, in particular to a compensation coefficient calculation method and device for a satellite reverse flywheel, equipment and a medium. The compensation coefficient range corresponding to the target angular momentum needing to be compensated for for the reverse flywheel is determined based on the target angular momentum needing to be compensated for for the reverse flywheel of the satellite, the preset variation corresponding to the target angular momentum, the inertia of the reverse flywheel and the maximum angular velocity of the rotating motor of the reverse flywheel; and calculating a residual angular momentum obtained after the compensation coefficient compensates the target angular momentum, and determining the compensation coefficient when the residual angular momentum is smaller than a preset minimum residual momentum as a target compensation coefficient of the reverse flywheel, so that accurate adjustment and calculation of the compensation coefficient of the satellite reverse flywheel can be realized. The accuracy of the compensation coefficient of the satellite reverse flywheel is improved, and then the shooting precision and stability of the rotary shooting device on the satellite are improved.

The invention relates to an attitude control parameter calculation method for a satellite with a large-flexibility solar panel. The attitude control parameter calculation method comprises the following steps: determining a target angle of rotation of a B axis of a two-dimensional solar panel driving mechanism relative to a zero position during sun tracking rotation of the solar panel; calculating a target rotational inertia by using the target angle and the satellite rotational inertia of the B axis of the two-dimensional solar panel driving mechanism at the zero position; and calculating PID control parameters by using the target rotational inertia. According to the attitude control parameter calculation method, for a satellite with a large-flexibility solar panel, self-adaptive calculation is carried out on the inertia change of the satellite when the B axis of the two-dimensional solar panel driving mechanism rotates to obtain the target rotational inertia, then the PID control parameters are calculated according to the target rotational inertia, and the robustness requirement of attitude control can be met.

A satellite control system fault detection method based on cooperative monitoring of energy and angular momentum comprises the following steps: firstly, setting a detection statistical period value, calculating a satellite average energy value, a theoretical angular momentum change value and a measurement angular momentum change value of each statistical period, then comparing the satellite average energy value, the theoretical angular momentum change value and the measurement angular momentum change value with a satellite average energy value of a previous statistical period, and scoring; comparing and grading the consistency of the theoretical angular momentum change value and the measured angular momentum change value, comprehensively judging the grading value, giving a satellite attitude control abnormity mark, and when the satellite attitude control abnormity is confirmed, performing subsequent processing to enable the satellite to autonomously switch into an attitude control safety capture mode.

The application relates to a flywheel speed control method and device based on a polynomial trajectory planning, belongs to the flywheel speed control method and device based on the polynomial trajectory planning, and solves the problem of how to reduce the impact disturbance introduced when a flywheel tracks an expected speed instruction of attitude control and improve the attitude stability of a satellite. The method comprises the following steps: receiving an expected speed instruction sent by an attitude control system; for each attitude control period of the attitude control system, a speed response curve between two adjacent expected speed instructions is constructed by adopting an expected speed sequence based on the polynomial trajectory planning, so that the acceleration is smooth and the jerk is continuous. The flywheel speed control method and device based on the polynomial trajectory planning are suitable for flywheel speed control in satellite attitude control.

The invention discloses a reaction wheel momentum management method for satellite redundancy configuration, and belongs to the technical field of satellite attitude control. According to the method, for reaction wheels in three-positive-one-oblique, pyramid or full-oblique redundant configuration, the core is that the characteristic that output torque of the configured reaction wheels is synthesized to be zero is utilized, the reaction wheel with the maximum difference value between the current rotating speed and the target rotating speed is recognized in real time through the preset target rotating speed to serve as a reference datum, and the reference datum serves as the guide; under the constraint condition that the resultant torque generated by all the reaction wheels is strictly guaranteed to be zero, the rotating speed output quantity of each reaction wheel is obtained, finally, the rotating speed of the corresponding reaction wheel at the current moment is adjusted based on the rotating speed output quantity of each reaction wheel, and momentum management of each reaction wheel in redundant configuration is completed. The method has many advantages in the aspects of momentum management, moment allowance control, control quality improvement, equipment service life prolonging and whole satellite energy saving of the satellite.

The invention relates to the technical field of space satellite control, in particular to a satellite attitude control method based on the Coulomb force control technology, electrode balls are configured on a satellite to enable the satellite to be electrified, and rotation starting and rotation eliminating actions of the satellite are achieved through polarity conversion between the electrode balls. Meanwhile, the problem that the satellite deviates due to coulomb force control is solved, control over the position, the attitude and the distance of the satellite is achieved, and support is provided for the follow-up space satellite control technology.

A satellite attitude preset performance control method based on control intensity quantification comprises the following steps: firstly, constructing a satellite attitude dynamic model containing a flexible mode to obtain an attitude tracking error dynamic model by aiming at a control intensity and control performance presetting problem according to a preset performance attitude control requirement under the limited control capability of a satellite; secondly, designing a preset performance controller by utilizing a preset performance function in a hyperbolic tangent form, and realizing satellite attitude control under preset performance; and finally, quantizing the control intensity under given controller parameters and a specified performance function to obtain an amplitude and a change speed range output by the controller. The control method designed by the invention has the characteristics that the realization is simple and convenient, the control intensity can be quantified, and the satellite attitude control performance can be preset.

This invention provides a novel integrated attitude sensor for satellites, relating to the fields of onboard attitude control and electronics. It includes: a star sensor body; a solar sensor body mounted on the star sensor body; a magnetometer body within the star sensor body; and a gyroscope body within the star sensor body. By integrating traditional discrete structures into a single device, the size and weight are reduced, and costs are significantly lowered. Magnetometer and gyroscope data are fused in real-time within an FPGA. This device features rapid adaptability, modular design supporting plug-and-play functionality, and integrated design improving EMC performance. It solves the problems of existing traditional satellite attitude control systems, where various attitude sensors are independently configured and equipped with dedicated processors, resulting in high hardware redundancy, large overall size and weight, high power consumption and cost, hindering further improvements in attitude measurement accuracy and dynamic performance, and leading to high overall design complexity.

The present application relates to the field of aerospace technology, especially to a kind of earth observation satellite attitude controller parameter self-tuning method, the scheme includes: obtaining the desired direction and actual direction of satellite, and determining the control torque of flywheel, the satellite is earth observation satellite;Based on the control torque, obtain initial PID parameter, initial PID parameter is PID controller simulation parameter;Based on the PID controller simulation method of flywheel physical constraint, construct the fitness function of PID controller simulation parameter based on different satellite imaging mode;Satellite attitude control is simulated by PID controller simulation parameter, based on fitness function and initial PID parameter, the optimal fitness corresponding to target PID parameter under each satellite imaging mode is determined by iterative updating through particle swarm algorithm, the optimal solution of PID parameter is target PID parameter by using particle swarm algorithm combined with fitness function for iterative updating, the accuracy and efficiency of parameter setting are improved.

A multimodal quantum federated learning system for satellites includes satellites and a central server configured to communicate with the satellites. The satellites include two or more of a bit-based satellite, a quantum-based satellite, and a hybrid satellite. The central server trains initial models including a bit-based network model and a quantum-based network model using training data, distributes the bit-based network model among the trained initial models to each of the bit-based satellite and the hybrid satellite, and distributes the quantum-based network model among the trained initial models to each of the quantum-based satellite and the hybrid 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.

This invention proposes an agile optical satellite linear array imaging method on the move, solving the problem of failing to obtain high-fidelity remote sensing images due to improper attitude control planning methods and imaging parameter selection and settings. The method includes: acquiring the satellite imaging task and selecting the satellite imaging mode; setting the evaluation benchmarks and their weights for each indicator in the imaging task; establishing an imaging parameter selection set, selecting one combination of imaging parameters for orbit calculation and attitude planning, calculating the imaging quality by correcting the yaw angle with the detector center point as a reference, and calculating the number of detector elements whose transfer functions meet the standards; calculating the comprehensive imaging evaluation parameter value of the imaging parameter combination based on the imaging quality value and the number of detector elements, using the set evaluation benchmarks and weights, and according to the comprehensive imaging evaluation function g; selecting the highest comprehensive evaluation parameter value from various imaging parameter combinations, and inputting the parameters of this combination to the satellite to execute the imaging task.

The invention belongs to the technical field of satellite attitude control, and relates to a distributed fault-tolerant attitude and vibration control method for a large flexible spacecraft, which comprises the following steps of: establishing a rigid-flexible coupling dynamic model of the large flexible spacecraft with a distributed actuator by adopting a Lagrange method, and taking an attitude control subsystem as a leader layer; the vibration suppression subsystem and the fault-tolerant control subsystem establish a layered Stackelberg game control framework for the follower layer, meanwhile, the performance functions of the three subsystems are defined, and strategy collaboration and dynamic equilibrium between at least two control targets are achieved through the game of spacecraft distributed fault-tolerant attitude and vibration control; according to the method, three-axis attitude stability and low-frequency vibration suppression of the target spacecraft are realized, and higher fault-tolerant capability is shown under the condition of node failure.

The invention discloses a sub-satellite on-orbit release and imaging observation method and related equipment, and relates to the technical field of satellite overall task planning and satellite attitude control, and the method comprises the steps: planning a release moment of releasing a sub-satellite when a main satellite is in a transit period; determining an orbit forecast result of the child satellite, and adjusting the attitude of the primary satellite to enable a visual axis of a release monitoring camera arranged on the primary satellite to be aligned with the child satellite; target identification is carried out based on image data collected by the release monitoring camera, whether a child satellite is captured is judged, and a first judgment result is obtained; and when the first judgment result is yes, controlling to release the monitoring camera to carry out imaging observation on the child satellite, and downloading image data through the ground data transmission station. According to the invention, the primary satellite can continuously carry out high-quality imaging observation on the secondary satellite, and meanwhile, the data transmission antenna is kept pointing to the ground, so that the real-time or quasi-real-time downloading of the image data through the ground data transmission station is ensured.

The present invention relates to a system and method for a ground-based simulation of an on-orbit operating environment of a laser communication terminal. The system comprises a plurality of subsystems, each subsystem comprising: a control unit, which performs data interaction with a timing and synchronization unit, a two-axis gimbal unit, a corresponding laser communication terminal, and a test unit; the timing and synchronization unit is configured for achieving time synchronization to a common reference between the respective subsystem and another subsystem; the two-axis gimbal unit is configured for simulating, on the basis of a satellite attitude control signal, an attitude of a satellite platform carrying a laser communication terminal, the laser communication terminal driving, on the basis of a pointing control signal of the control unit, a rotating mechanism to adjust a pose of the laser communication terminal; the test unit is configured for testing the laser communication terminal; and an optical path assembly unit configured for forming or disconnecting a free-space laser link with an optical path assembly unit of another subsystem. The present invention enables simulation of an on-orbit operating environment of a laser communication terminal, and performs comprehensive and dynamic functional testing for laser communication terminals.

The application discloses a satellite attitude super-agile wobble scanning neural network minimum parameter learning sliding mode control method and belongs to the technical field of micro-satellite attitude control. The application discloses a satellite attitude super-agile wobble scanning neural network minimum parameter learning sliding mode control method and belongs to the technical field of micro-satellite attitude control. The method is used for converging the system state to zero within a limited time through the fast terminal sliding mode control, and the switching term in the sliding mode is approximated through the radial basis function neural network, so that the purpose of weakening the chattering is achieved. Meanwhile, the minimum parameter learning method is designed to replace the weight value adjustment of the radial basis function neural network, the weight value matrix is prevented from being operated in a large amount, the controller efficiency is effectively improved, and the method has the characteristics of good convergence, imaging and control effect.

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.