2040 results about "Spacecraft" patented technology

The invention belongs to the technical field of spacecraft pose estimation, and particularly relates to a cross-domain spacecraft pose estimation method based on mask self-distillation domain adaptation, and the method comprises the steps: 1, inputting a complete image, and employing a Faster R-CNN algorithm to position a spacecraft bounding box; the robustness of the model is improved by applying a track environment data enhancement strategy; and extracting a target ROI region as key point regression network input based on the detection frame. 2, dividing ROI (Region of Interest) data of a source domain and a target domain; and optimizing heat map supervision loss learning key point positioning knowledge. 3, inputting random mask enhanced target domain data into the student model; inputting original target domain data into the teacher model; a learnable shared prototype space is constructed, and self-distillation is guided through heat map consistency loss and semantic consistency loss. And 4, jointly optimizing the loss of the key point regression network 3, and realizing progressive migration of source domain annotation knowledge to a target domain. And 5, solving the 6D pose of the spacecraft relative to the camera through the EPnP. According to the invention, robust six-degree-of-freedom pose estimation of the target spacecraft is realized.

The invention discloses a cabin butt joint guide rail compensation method based on point cloud, and relates to the technical field of spacecraft cabin automatic butt joint, and the method comprises the following steps: building a system coordinate system, and carrying out initial calibration; carrying out guide rail error modeling; carrying out six-degree-of-freedom platform error correction modeling; point cloud data acquisition and pose measurement; calculating error compensation control; adjusting and closed-loop feedback are executed; according to the cabin butt joint guide rail compensation method based on the point cloud, modeling and real-time compensation control are carried out on pose errors in the cabin butt joint process under guide of the guide rail by fusing the three-dimensional point cloud sensing technology and six-degree-of-freedom platform control; the problems that in the prior art, guide rail errors cannot be automatically corrected, and the butt joint precision is insufficient are solved, and accurate butt joint, safety and reliability of the large cabin section in the automatic assembling process are guaranteed.

The invention discloses an intelligent fusion analysis and decision support system and method for multi-source heterogeneous data of a spacecraft. The intelligent fusion analysis and decision support system for the multi-source heterogeneous data of the spacecraft and the intelligent fusion analysis and decision support method for the multi-source heterogeneous data of the spacecraft are included. The system comprises a dynamic adaptation access layer, a unified semantic modeling layer, a space-time semantic fusion engine, a digital twin collaborative decision-making center, a knowledge enhancement decision-making auxiliary subsystem, a cross-department collaborative interaction platform and a side cloud collaborative intelligent processing framework. The method comprises the steps of multi-source heterogeneous information dynamic access and semantic modeling, space-time and semantic deep fusion, digital twin driven decision deduction, knowledge enhanced decision assistance, cross-department collaborative decision and interaction, edge cloud collaborative intelligent processing and the like. According to the method, the safety, reliability and decision-making efficiency of spacecraft tasks can be remarkably improved, and the method is suitable for development, testing and on-orbit task stages of various spacecrafts.

The invention discloses an on-orbit spacecraft attitude estimation method and system based on ISAR image feature selection, and belongs to the field of aerospace control systems. The method comprises the following steps: acquiring an ISAR image of a spacecraft, and acquiring a complex linear structure set and three-dimensional feature points of an on-orbit spacecraft; then, according to the obtained three-dimensional-two-dimensional projection model, the CRLB of each reference structure in the complex linear structure set is deduced to carry out attitude estimation error analysis; calculating the trace of the CRLB covariance matrix of each reference structure to select an optimal feature structure, correcting the scattering point trace of the optimal feature structure by using polynomial fitting, and switching the reference structures as a new optimal feature structure according to the scattering point loss rate and a preset sequence; and optimizing the spacecraft attitude angle solving function by using a particle swarm and LM hybrid algorithm to obtain attitude angle parameters. The target with high-precision target attitude real-time estimation can be completed aiming at the problems that high-order frequency change in a dynamic environment is difficult to capture and resolution and noise suppression are contradictory due to a fixed window time-frequency analysis method.

Systems and methods for autonomous space-mission coordination integrate ground-based prediction with on-orbit execution. A processing system in a ground station may ingest multi-source environmental data, produce near-term nowcasts through an prediction model, rank pending spacecraft tasks against current resource telemetry, select a high-value task subset, convert the subset into time-tagged command packets, and transmit the packets through a communications link. A processing system aboard each spacecraft may receive the packets, merge them into a persistent schedule, and at each time tag slews attitude, activate an imaging or radar sensor with specified parameters, capture data, and store the data in non-volatile memory. The spacecraft may generate quality metrics for the captured data and return the metrics to the ground station.

The invention relates to a Bayesian KAN fused aircraft rapid multidisciplinary uncertainty analysis method, and belongs to the technical field of spacecraft manufacturing and application. According to the method, Bayesian learning is combined with a Colmogorov-Arnodel network, a Bayesian KAN neural network which can cope with high-dimensional input uncertainty, can quantify model cognitive uncertainty and is higher in fitting capacity is provided, a BKAN network agent model is constructed for a subsubject analysis model, time-consuming numerical simulation of each subsubject is replaced, and the calculation amount is reduced; with reference to a basic framework of a global sensitivity equation GSE of a deterministic field coupling multidisciplinary system, input random input and model cognition uncertainty are introduced, and expressions of mean values and covariances of disciplinary coupling variables and system responses are deduced on the GSE framework; a rapid semi-analytical multidisciplinary random and model cognition hybrid uncertainty propagation method is established, efficient multidisciplinary uncertainty analysis is achieved, and the problems that an existing method is poor in precision and large in calculation amount are solved.

The invention discloses a reinforcement learning-based cluster spacecraft multi-target intelligent cooperative tracking method. The method comprises the steps of establishing a relative motion relationship between a tracking spacecraft and a target spacecraft through a nonlinear orbit kinetic equation; designing a double-layer game framework, constructing a non-zero sum game in the cluster to drive cooperation, and introducing a zero and maximum and minimum game between the cluster and a target to describe a confrontation relationship; based on communication topology and relative state information, defining a local error term and a performance index function of the tracking spacecraft and the target spacecraft, and defining an optimization target of each agent under cooperation and confrontation; and iteratively updating the value function and the control law. According to the invention, collaborative tracking can be realized without defining a system model. The method can be widely applied to the field of cluster control.

The invention discloses a high-heat-flux heat dissipation system for an on-orbit computing power cluster, and belongs to the technical field of spacecraft heat control. The system is independent of a satellite conventional thermal control platform and is specially used for discharging waste heat of an AI calculation array. The system comprises a heat capture unit, a pumpless heat transfer unit and a foldable heat dissipation unit. The heat transmission unit comprises a micro-channel flexible pulsating heat pipe (such as a diamond film material), and the heat pipe spans a folding area and has the functions of structural connection and heat transmission, so that a thermal low-impedance independent channel relative to a satellite structure is formed. Through the design of removing a mechanical pump and removing a mechanical hinge, the problems of large weight, micro-vibration interference, high contact thermal resistance and the like of a heat dissipation system of the computing power satellite are solved, and preferential directional heat dissipation of high-flux heat is realized.

The invention discloses a high-strength heat-insulation aerogel structure with a dual multi-scale design and a preparation method thereof, relates to the technical field of thermal protection of spacecrafts, and in particular relates to the high-strength heat-insulation aerogel structure with the dual multi-scale design and the preparation method thereof. The invention aims to solve the problems of insufficient structural rigidity and poorer thermal protection property and thermal stability of the traditional spacecraft surface aerogel structure under the conditions of extreme high temperature and mechanical load. The structure is composed of an upper carbon-carbon composite material plate, a lower carbon-carbon composite material plate, an alloy skeleton and a filler. The method comprises the following steps: 1, preparing a multi-scale C / C composite material; 2, designing and preparing a titanium alloy honeycomb framework; 3, proportioning the multi-scale aerogel; 4, in-mold in-situ gelation; 5, drying and heat treatment; and 6, assembling and solidifying. The porosity of the aerogel is 85-98%, the structure has high rigidity and thermal protection performance, and the aerogel is applied to the field of aerospace thermal protection.

The invention relates to a near-earth spacecraft strong transverse maneuvering reentry trajectory planning method. The method comprises the following steps: determining basic parameters of return reentry trajectory planning; calculating a reentry trajectory of the maneuver without transverse offset; distributing a transverse maneuver amount, and establishing a transverse maneuver planning equation; calculating a braking yaw angle; calculating an initial value of the heeling angle offset; establishing a funnel guidance law under the condition of transverse offset maneuvering; according to the determined return and reentry trajectory planning basic parameters, the total transverse offset maneuver amount, the braking yaw angle and the funnel guidance law under the transverse offset maneuver condition, the initial value of the heeling angle offset amount, the time of the braking point relative to the return circle ascending node and the braking starting duration serve as initial values, reentry trajectory parameters are iteratively solved, and reentry trajectory planning is completed. According to the method, the problem of rapid planning of the reentry trajectory when the ground proximity device returns to the designated position under the condition of large initial transverse deviation is solved, and meanwhile, the problem of rapid planning of the reentry trajectory when the landing point needs to avoid the designated area can be solved.

The application provides a spacecraft thin-wall cabin lightening groove online measurement and compensation processing method, comprising the following steps: installing a thickness measuring head; setting a measurement point on the surface of each to-be-processed lightening groove and obtaining a theoretical thickness and an outer contour position; measuring the actual thickness of each measurement point and obtaining an actual outer contour position; calculating a compensation value of each measurement point according to the theoretical value and the measured value of the thickness and the outer contour; judging whether the compensation requirement is met according to the compensation value, if yes, processing after compensating the tool length based on the comprehensive compensation value, if not, reconstructing the processing surface based on the actual thickness and the outer contour position, and processing by using the initial tool length based on the reconstructed processing surface. The application calculates the compensation value according to the contour and thickness data, and finally processes by calling the compensation value, thereby improving the automation degree of measurement and processing, realizing efficient and high-precision processing of the spacecraft thin-wall cabin lightening groove and ensuring the thickness size precision.

The invention relates to an earth-moon space low-thrust spacecraft fuel optimal collision avoidance trajectory implementation method, which comprises the following steps of: constructing a kinetic model based on calculation parameters of a spacecraft collision avoidance problem, and establishing a fuel optimal performance index of the spacecraft collision avoidance problem through a homotopy method; the optimal control problem is converted into a two-point boundary value problem, targeting solving is carried out, and the state and co-state parameters of the optimal orbit are obtained; and according to the obtained state and co-state parameters of the optimal orbit, solving to obtain collision avoidance optimal control and collision probability of the spacecraft. According to the method, the fuel optimal control problem is solved by adopting the indirect method, the convergence difficulty of the low-thrust fuel optimal problem is reduced through the homotopy method, and when the thrust is calculated, the calculation efficiency is high, the precision is high, and the method is suitable for the fuel optimal collision avoidance scene of the low-thrust spacecraft for a single target in the earth-moon space.

Combined aerospace system is a multi-stage technology to deliver payloads to orbit and to midair. The system comprises a carrier aircraft, a tow aircraft, a towed spaceplane or a rocket plane or a glider. The carrier aircraft, coupled with the spaceplane and / or the rocket plane or several gliders, takeoff from a runway and ascends to a staging altitude and speed under the carrier aircraft power. After staging, the tow aircraft will pick up the spaceplane or the rocket plane or the glider from the carrier aircraft in flight. The towed spaceplane or the rocket plane or the glider continues ascending and accelerating by the tow aircraft. The spaceplane or the rocket plane will disconnect from the tow aircraft and continue to ascend and accelerate in an atmosphere by using the ramjet engine. The spaceplane or the rocket plane equipped with a propulsion unit is used to place the spacecraft into the earth orbit. The glider will disconnect from the tow aircraft and lands in designated place. Gliders can be used for instantly delivering of correspondence, parcels, and civil goods from a supplier to wide variety of individual consumers or corporate customers.

The invention relates to the technical field of space operation equipment, in particular to an electric tool applied to on-orbit operation of a space mechanical arm. The high-precision control and the autonomous operation capability of the tail end of the space manipulator on the tool state in a complex space environment are realized, and the reliability, the flexibility and the automation degree of on-orbit operation are greatly improved. Compared with the prior art, the electric tool has the advantages that the structural integration degree, the function expandability and the intelligent feedback control are remarkably improved, and the electric tool is innovatively designed in the aspects of automatic storage of the bit, operation continuity and space environment adaptability; and the execution efficiency and the safety level of on-orbit maintenance and construction tasks of space stations, satellites and other spacecrafts can be effectively improved.

The locking and releasing device for the spacecraft comprises a base, a shell, a sliding sleeve, a pin puller, a split nut assembly and a first spring, the base is of a cylindrical structure, and a first mounting plate and a second mounting plate are fixed to the two axial ends of the base correspondingly; the sliding sleeve slidably sleeves the base and is connected with the first mounting plate through the first spring. The shell sleeves the sliding sleeve at an interval and is fixed to the first mounting plate and the second mounting plate. The pin puller is mounted on the shell, and one end of the pin puller extends into the shell; a plurality of notches evenly distributed in the circumferential direction are formed in the side wall of the base, a first through hole communicated with the interior of the base and used for ejecting the locking rod is formed in the middle of the first mounting plate, a split nut assembly used for locking the locking rod is mounted in the base, and a plurality of sets of connecting lug plates evenly distributed in the circumferential direction are fixed to the outer side wall of the split nut assembly. The multiple sets of connecting lug plates are arranged in one-to-one correspondence with the multiple notches and penetrate through the notches through connecting rods to be hinged to the sliding sleeve.

The invention provides a spacecraft thermal test thermal response automatic test method and system. The method comprises the steps that S1, preset configuration information is acquired, and a resistance interpretation test is carried out on a current power output loop; s2, initial test current is applied to the power output loop passing the test, and the temperatures of all test points are obtained regularly; and S3, judging whether the temperature change of the measuring point meets the response condition or not, if so, judging that the power output loop is matched with the measuring point meeting the response condition, and testing the next power output loop, and if not, returning to the step S2 and increasing the test current step by step until the measuring point meeting the matching condition exists. The testing process is high in safety, it is initially determined that the configuration of the power output loop is not abnormal, then the matching performance of the temperature measuring point and the power output loop is determined step by step in a stepped power-up mode, and the problems that due to the fact that the current applied to the power output loop is too large for a long time, a heating piece is burnt or the temperature of a product is abnormal are completely eradicated.

The invention discloses a dual-path non-cooperative spacecraft pose estimation method based on uncertainty perception, and belongs to the technical field of spacecraft pose estimation. According to the method, a dual-path prediction framework based on a shared backbone network is constructed and comprises a geometric reasoning path and a global context sensing path. In the training stage, prediction results of the two paths are aligned through a geometric consistency loss function, and joint optimization is carried out by combining multi-task losses such as key point regression, classification, rotation and translation regression and the like. In the inference stage, the pose estimation result is adaptively fused according to the predicted variance output by each path, when the variance is low, the output of the global context sensing path is directly adopted, otherwise, the key point projections of the two paths are combined, and the final pose is recovered. The method has higher estimation precision, robustness and interpretability in complex space environments such as shielding, illumination variation and large-scale variation, and provides reliable visual navigation support for on-orbit autonomous tasks.

The invention relates to a step-by-step winding and unfolding ultra-large flexible solar cell array and an unfolding method, and belongs to the technical field of space solar cell arrays. The battery array comprises a battery array body structure, a pressing and releasing mechanism, an unfolding and locking mechanism and an extension arm mechanism, the cell array body structure has a folded state and an unfolded state; under folding, pressing force is provided through the pressing and releasing mechanism, so that the device is folded on the side walls of the two sides of the satellite body and can bear the mechanical load of the launching section; during unfolding, under the action of the unfolding locking mechanism, the folding mechanism and the stretching arm mechanism, the folding mechanism is unfolded in a distributed manner according to a time sequence; and after being unfolded, the antenna can rotate by 360 degrees to provide energy for satellites. The invention has the advantages of small folding volume, large unfolding area and light weight, and solves the contradiction that the power demand of a spacecraft power supply is large and the envelope size of a launching section is small.

The invention discloses a method for predicting the remaining service life of a control moment gyroscope based on parameter fusion, and relates to the technical field of health management of spacecraft attitude control systems. According to the method, optimal fusion of measurement parameters and implicit parameters is realized through adversarial learning feature extraction and a channel attention mechanism, and the method is particularly suitable for high-precision residual life prediction of the spacecraft control moment gyroscope under variable working conditions. The method comprises four main steps of data preparation and multi-source input, adversarial learning feature extraction, SE-Attention feature fusion and three-stage training optimization. In the data preparation step, original time sequences of voltage and current of a rotor motor are collected through a built-in electrical sensor of a CMG, and key physical parameters of a full-film lubrication friction coefficient, lubricating oil viscosity, a bearing clearance and contact stiffness are inverted based on an electromechanical coupling model; in the adversarial learning step, a game framework of a feature extractor and a working condition discriminator is constructed, and working condition-independent robust feature extraction is realized through a gradient inversion layer; in the SE-Attention feature fusion step, importance weights are adaptively distributed to different feature channels through an extrusion-excitation mechanism; and the three-stage training optimization step adopts a preheating-confrontation-fine tuning strategy to ensure the network convergence. The method solves the problem of domain adaptation, has the technical advantages of strong robustness, high precision, strong generalization, interpretability and self-adaptation, and can realize high-precision RUL prediction under variable working conditions.

The invention discloses a light-weight cabin efficient self-adaption machining method based on cutting parameter learning, solves the cutting parameter self-adaption problem in light-weight cabin machining, and belongs to the technical field of spacecraft structural part manufacturing or numerical control self-adaption machining. The method aims at improving the machining quality and the machining efficiency, minimizing the cutting force and maximizing the material removal rate, a mathematical model of two targets and cutting parameters is established, an orthogonal process test is designed to analyze the response characteristics of the two targets to the cutting parameters, the influence of the cutting parameters to the cutting force in the actual cutting process is analyzed, and the machining precision is improved. Optimized operation data and a tool database are formed; and learning of adaptive parameters is realized by applying an ACM adaptive optimization technology, and adaptive processing is carried out.

A measurement and control simulation device based on a measurement and control data transmission integrated system comprises an upper computer display control test module, a PC host module, a signal processing master controller module, an isolation DC / DC module, a down-conversion module, a filtering module, an isolator module and an outer end face interface module. The system has the characteristics of multiple modulation systems, a spread spectrum function, controllable output power, Doppler frequency change and the like, can be used for simulating a star chain control system link test in a ground test system, and fully improves the index assessment performance during system docking; meanwhile, the design of a down-conversion link meets the requirements for the strength and sensitivity of the received signal, the stability and accuracy of the radio-frequency signal after frequency conversion in a low-power injection state are remarkably improved, and a new scheme is provided for a high-frequency conversion mode. Besides, the display control module of the equipment has the functions of multiplexing, reconstruction, multi-data interaction and the like, so that multifunctional and multi-task full-system test coverage can be realized, the cost investment of test systems such as spacecrafts and the like is greatly reduced, and the multifunctional index test coverage efficiency is improved.

The invention provides a satellite power supply system topology architecture and a maximum power point tracking control method thereof, and aims to improve the energy utilization efficiency of a solar cell array of a spinning body spacecraft under a complex illumination condition. The architecture adopts a combined topology architecture of a preceding-stage sequence switch maximum power point shunt regulation (S3MPR) circuit and a backward-stage four-switch Buck-Boost (FSBB) converter, and is suitable for a special operation environment of a spinning body spacecraft power supply system. The system adopts a bus full adjustment mode, enables the DC bus voltage to be stabilized near the maximum power point voltage of the photovoltaic array through a double closed-loop control strategy, and carries out real-time shunting adjustment on an S3MPR circuit in combination with a multi-peak maximum power point tracking (MPPT) algorithm, thereby ensuring the efficient operation of the photovoltaic array. According to the invention, high-efficiency energy tracking and stable charging can be realized under a complex illumination condition, and the reliability and the energy utilization rate of a spinning body spaceflight power supply system are remarkably improved.

The invention discloses a non-cooperative spacecraft monocular attitude estimation method and system. The method comprises the following steps: acquiring a shot image of a non-cooperative spacecraft; inputting the shot image into a pre-constructed target attitude estimation model to obtain attitude estimation of the non-cooperative spacecraft, including a position vector and a direction vector; wherein the training process of the model specifically comprises the steps of fusing a scale weight and a space weight, and estimating a feature map subjected to weighted fusion through a position estimation head to obtain an estimated position vector; estimating the weighted local feature map through a direction estimation head to obtain an estimated direction vector; a weight adjustment amount is dynamically generated, the fusion weight is optimized and updated, and attitude estimation is carried out again; and carrying out iterative optimization on the model to obtain an optimal target attitude estimation model. According to the method, a relatively low attitude estimation error can be kept while the calculation speed and the calculation frequency are increased.

A tiered spacecraft docking station is adapted to facilitate docking of spacecraft within outer space. A first tier includes a first frame enclosing a first area. A first net-like mesh is coupled to the first frame and fills the first area enclosed by the first frame. A second tier includes a second frame enclosing a second area. A second net-like mesh is coupled to the second frame and fills the second area enclosed by the second frame. A plurality of support beams attach the first frame to the second frame. A lander is used to slow and stop a spacecraft on a celestial body. The lander includes a first and a second webbed structure and a decelerator coupled to the first webbed structure and / or the second webbed structure. The decelerator maintains a tension in the first and / or second webbed structure below a predetermined threshold.

The invention discloses a spacecraft component identification method based on a receptive field enhanced network, and relates to the technical field of remote sensing, the network is composed of a receptive field enhanced feature extraction encoder and a double-pooling window self-attention feature fusion decoder, and the encoder is composed of a backbone network ConvNeXt-B and a receptive field enhanced feature extraction module. The receptive field enhanced feature extraction module adopts asymmetric convolution and multi-scale receptive field design so as to enhance the multi-stage feature capture capability of a small-size component and relieve the feature loss problem. The decoder is composed of a double-pooling window self-attention feature fusion module, and the double-pooling window self-attention feature fusion module comprises a window self-attention mechanism and a double-pooling feature fusion module. And the double-pooling feature fusion module is used for fusing features through double-pooling operation and deep convolution, so that the detection precision of the small target is remarkably improved.

An apparatus for transmitting heat from a component side to a radiator side on a spacecraft includes a component side that contains a heater, an electrical component, an optional first conduction link, and a first portion of a switchable thermal link and a radiator side that includes a radiator, an optional second conduction link and a second portion of a switchable thermal link. The first portion of the switchable thermal link and the second portion of the switchable link are in operative thermal communication with the component side and the radiator side respectively. The first portion of the switchable thermal link and the second portion of the switchable thermal link are reversibly attachable and detachable. The switchable thermal link is operative to receive and radiate thermal energy from the component side to the radiator side.

The invention discloses a spacecraft Hall propulsion device test working medium flow stability control system, which comprises a sensing module for detecting working condition state parameters of working medium supply equipment, and an adaptation module for exploring current working condition control parameters by combining detection data of the sensing module and similar working condition control parameter empirical data fusion, the adjusting module is used for receiving the control parameters and dynamically adjusting the adjusting weight of the adjusting device according to different phase change periods, an experience library unit and a self-adaption unit are arranged in the adaption module, the experience library unit provides experience data of similar working conditions and control parameters, and the self-adaption unit provides self-adaption data of the control parameters. And the autonomous adaptation unit fuses and explores control parameters by constructing a neural network adaptation model, so that working condition automatic identification and control parameter autonomous adaptation under the environment that the working medium type or the working medium state rapidly changes are realized; the problems of low control precision and slow response speed caused by manual presetting of control parameters of different working media in a traditional control technology are solved, and stable control of the working medium flow is realized.

The invention discloses a large-size spacecraft dynamic six-degree-of-freedom pose high-precision measurement method, solves the problems of shielding failure, low measurement precision and the like in large-size spacecraft dynamic tail end target support measurement, belongs to the technical field of spacecraft dynamic measurement, and is mainly suitable for large-size spacecraft tail end antenna dynamic working conditions. For example, the three laser trackers are adopted for cooperative work, a synchronous trigger sends a unified trigger signal to all the laser trackers, and it is ensured that the three trackers collect three-dimensional coordinate data of a target point at the strict same moment. And finally realizing six-degree-of-freedom pose high-precision measurement of the large-size spacecraft motion tail end target bracket by combining a tracker space station planning strategy of a non-coplanar optimization algorithm of a minimum position precision attenuation factor and utilizing a dynamic six-degree-of-freedom pose resolving algorithm fused with measurement data based on a space target coordinate obtained by synchronous triggering.

The invention discloses a camera imaging control method and device for midnight sunlight avoidance, and belongs to the field of spacecraft attitude orbit control. The method comprises the steps of judging whether a satellite is currently in a midnight sunlight avoidance stage; when the satellite is not in the midnight sunlight avoidance stage, if the midnight sunlight avoidance condition is met, initializing to enter the midnight sunlight avoidance stage, and setting the satellite to quickly maneuver to a preset target avoidance trajectory; when the satellite is in the midnight sunlight avoidance stage, if the satellite receives an imaging instruction meeting the requirement, the satellite is maneuvered to an imaging point for imaging, and otherwise, the satellite is maintained to run on a target avoidance track; and when the satellite is in the imaging state of the midnight sunshine avoidance stage, if the satellite is in a sunshine receiving state, quickly maneuvering the satellite to the target avoidance track, otherwise, maintaining the imaging state of the satellite. According to the invention, the imaging safety of the camera during the midnight avoidance period can be realized, and the use efficiency of the camera load and the user experience are improved.

The invention relates to a ground test gravity unloading device for simulating a space microgravity environment. A base is fixedly arranged at the bottom of a fixed frame; a suspension system is arranged above the base and comprises a fixed pulley block mounted on the fixed frame, a pull rope is wound on the fixed pulley block, the bottom of one end of the pull rope is connected with a dynamometer through a tension spring, and the bottom of the other end of the pull rope is wound on a winch mounted on the base; the hoisting system comprises an external excitation device which is suspended right above the detected object; the displacement sensing system comprises a displacement sensor. The device is simple in principle and construction mode, has a relatively ideal controlled unloading mode, has the test advantage of controllable fine adjustment, is mainly used for simulating a space micro-vibration / microgravity environment to carry out a ground test, reproduces a space environment on the ground to the greatest extent so as to verify the functions and performance of a spacecraft and components thereof, and has a wide application prospect. For example, whether the vibration isolator can work reliably on orbit is verified, so that the on-orbit risk is remarkably reduced, the research and development cycle is shortened, and huge cost is saved.