2532 results about "Spacecraft" patented technology

The invention provides a spacecraft full-cycle monitoring and operation and maintenance decision-making system based on digital twinning, and belongs to the field of spacecraft full-life-cycle intelligent operation and maintenance. Structural data and environmental data are monitored in real time through a multi-source sensor network deployed on a spacecraft structure; fusing physical entity monitoring data and damage evolution rule virtual data in a twin database, constructing a structure damage resistance and mechanical property evaluation model, and executing geometric modeling, physical attribute analysis, behavior response simulation and damage evolution rule modeling; based on a simulation result of the virtual model, life prediction and reusability evaluation are carried out, and an operation and maintenance decision instruction is generated; and feeding back the operation and maintenance decision instruction to the physical entity to execute control, and updating state data in the twin database to perform iterative optimization. According to the method, full-scene coverage from design demonstration to scrap recovery is supported, the problem of stage splitting in the traditional technology is solved, and a closed-loop optimization mechanism is formed through real-time interaction between the physical entity and the virtual model.

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 provides a spacecraft quartz powder heat insulation tile which is prepared by adopting high-purity quartz powder (SiO2) as a main raw material and combining a high-temperature-resistant ceramic fiber reinforcement and a nano aerogel heat insulation layer through a high-temperature sintering process. The heat insulation tile has the following advantages: 1, the heat insulation tile has excellent high temperature resistance, and can tolerate 1600 DEG C high temperature for a long time and tolerate 1800 DEG C instantaneous thermal shock for a short time; 2, repeated use: the structure is stable, and high strength is still kept after multiple thermal cycles; light weight: the density is low (less than or equal to 0.5 g / cm < 3 >), and the load of a spacecraft is reduced; and 4, the thermal conductivity is low (less than or equal to 0.05 W / m.K), and heat transfer is effectively isolated. The method is suitable for thermal protection systems (TPS) of spacecrafts, such as thermal insulation of reentry capsules and rocket engine nozzles.

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 relates to the technical field of spacecrafts, and discloses a satellite solid storage and data transmission integrated system which comprises a data transmission solid storage master control FPGA, an FPGA loading and refreshing system and an ARM system, the FPGA loading and refreshing system and the ARM system are connected with the data transmission solid storage master control FPGA, the ARM system is connected with the FPGA loading and refreshing system, and meanwhile the ARM system is connected with satellite service through a CAN bus; the FPGA loads and refreshes a system, and periodically refreshes rewrite configuration or system reset rewrite configuration to ensure normal operation of the configuration; and the ARM system forwards the satellite remote control instruction received from the CAN bus to the data transmission solid memory master control FPGA and the loading FPGA through the SPI bus. According to the technical scheme, storage and transmission are efficiently integrated, the data storage rate, the data transmission rate and the data transmission capacity are improved, the system architecture is simplified, flexible selection of a data transmission modulation mode and the data transmission rate is achieved, the system reliability is improved, data access and transmission can be more accurately regulated and controlled through control and management of the FPGA, and the system reliability is improved. And the system instability caused by hardware faults or interface problems is effectively reduced.

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.

The invention provides a PPO-based spacecraft attitude and orbit integrated game control method and device, and the method comprises the steps: building a spacecraft attitude and orbit integrated model, and carrying out the modeling of the interaction process of a tracking spacecraft and a target spacecraft as a non-cooperative zero-sum game; setting a multi-agent deep reinforcement learning framework, wherein the tracking spacecraft and the target spacecraft are respectively configured with an independent strategy network and an independent value network; training is carried out in a double-agent PPO mode, control instructions of orbit thrust and attitude torque are generated through the strategy network, and the state value is evaluated through the value network; and controlling and guiding the behavior of the spacecraft according to the trained strategy. According to the method, the autonomous decision-making capability and cooperative control performance of the spacecraft in complex tasks are effectively improved, and the method is suitable for high-precision aerospace tasks such as interception of non-cooperative targets, autonomous trajectory planning and attitude control.

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 rigid spacecraft attitude control method and system based on finite time constraint, and belongs to the technical field of spacecraft attitude control. Comprising the following steps: respectively constructing a spacecraft attitude control system model and a finite time convergence performance function containing a preset time constant; according to the finite time convergence performance function, dynamic attitude constraint is applied to the system state of the spacecraft attitude control system model; converting the attitude constraint into differential homeomorphic mapping of an unconstrained space; obtaining an error conversion variable according to differential homeomorphic mapping; and designing a virtual control law and constructing an error compensation module according to the error conversion variable, and designing an actual control input law according to the virtual control law and the error compensation module. According to the method, rapid convergence of the spacecraft attitude within the preset time is ensured, the dynamic response precision of the spacecraft in the rapid attitude maneuver process is remarkably improved, the control period delay is effectively avoided, and the power consumption is reduced.

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 spacecraft complex service damage automatic identification method based on multi-modal data fusion, and the method comprises the steps: S1, obtaining a real infrared thermal image data set S1 or a real vibration data set S2 for different spacecraft complex service damage test pieces; s2, performing expansion processing on the S1 and the S2 to obtain an infrared data set Sir and a vibration data set Svib, and combining the Svib and the Sir to obtain a multi-modal data set S; s3, constructing a complex service damage identification model, wherein the complex service damage identification model takes the labeled multi-modal data set S as a training sample; and S5, fusing prediction results through a fusion decision module in the complex service damage identification model so as to complete comprehensive judgment of the complex service damage state. According to the spacecraft complex service damage automatic identification method based on multi-modal data fusion, the number of data sets is remarkably increased under the condition that test piece samples are limited, and the requirement of deep learning model training for large-scale data is met.

The embodiment of the invention provides a measurement and control link multiple access method and device based on a high-dimensional discrete space, electronic equipment and a storage medium. The method comprises the following steps: constructing the high-dimensional discrete space based on an initial orthogonal vector coordinate system; the high-dimensional discrete space comprises space resources of at least one of a time domain, a frequency domain and a code domain; in a high-dimensional discrete space, constructing a non-stipulated coordinate system matched with the control scene requirement by taking the initial stipulated coordinate system as a reference; constructing a dual mapping relation matrix according to a non-stipulated coordinate system of the initial stipulated coordinate system, and updating the dual mapping relation matrix based on the measured data; and accessing at least one satellite or spacecraft to the ground station based on the updated dual mapping relation matrix and the constellation measurement and control link, thereby realizing effective expansion of the measurement and control link multiple access technology, and improving the measurement and control link multiple access efficiency based on the available discrete space of the multi-dimensional joint variable.

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 satellite deck coupling connection mechanism, a satellite and a satellite deck coupling connection method, belongs to the technical field of satellite structures, and particularly relates to the satellite deck coupling connection mechanism and the satellite deck coupling connection method. The problems of low assembly efficiency, uncontrollable thermal stress and difficulty in on-orbit maintenance in the existing bolt connection technology are solved; the mechanism comprises interlocking units which are arranged on the butt joint edges of the two cabin plates in pairs. The satellite cabin plate coupling connection mechanism, the satellite and the satellite cabin plate coupling connection method are applied to modular cabin plates of spacecrafts such as the satellite and the like, and bolt-free interlocking connection aims at overcoming the defects of traditional bolt connection in the aspects of rapid disassembly and assembly, thermal deformation adaptability and the like.

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 method, system and device for scheduling tasks of low-orbit constellations for dynamic targets in space under a fully distributed satellite system, belonging to the field of on-orbit mission planning technology for spacecraft, and solving the problems of high technical cost, low flexibility and low stability of low-orbit constellation task scheduling. The method of the invention comprises: using an autonomous distributed task scheduling method based on CCGA to solve the multi-satellite scheduling problem, decomposing the multi-satellite scheduling problem into a single-satellite scheduling problem, and solving the optimal scheduling solution for each satellite separately to form an overall optimal scheduling solution. The scheduling process consists of local optimization and global negotiation. Each satellite independently explores the scheduling problem of the satellite for local optimization. At the same time, a ring topology is used to form an inter-satellite communication network for global negotiation within the constellation. The observation effect of the scheduling solution is comprehensively evaluated by the weighted sum of the satellite contribution, the number of switching times and the satellite load. The task benefit evaluation index is introduced to assist in the optimization of the single-satellite scheduling solution. The invention is suitable for low-orbit constellation task scheduling.

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.

The invention discloses a spaceflight electromagnetic relay failure early warning and diagnosis method based on multi-physics field coupling, relates to the technical field of spacecraft electrical system reliability monitoring, and is used for solving the problems of lack of comprehensive consideration of multi-physics field coupling characteristics under spaceflight special working conditions and low interference resistance to complex environments. The method comprises the following steps: monitoring a network state, obtaining real-time data, calculating an optimal transmission strategy by using an intelligent regulation and control model, dynamically adjusting a coding mode, adaptively adjusting a transmission rate, further ensuring data integrity through an error code detection and error correction mechanism, and adjusting transmission parameters according to a real-time network condition in combination with packet optimization and flow priority management. Effective recognition and early warning of typical failure modes such as contact adhesion and coil burnout are achieved, and the failure early warning accuracy and diagnosis efficiency of the spaceflight electromagnetic relay in complex environments such as high radiation, extreme temperature and strong vibration are improved.

The invention discloses a spacecraft sensor abnormal data recovery method based on spatio-temporal information fusion, and relates to a spacecraft sensor abnormal data recovery method. The method aims at solving the problems that in an existing spacecraft sensor abnormal data recovery technology, the signal-to-noise ratio is deteriorated due to environment interference, and multi-dimensional feature modeling has limitation and is insufficient in self-adaptive capacity. The method comprises the following steps: step 1, preprocessing original sensor data through a sliding mean filter; 2, constructing a time-space dual-scale feature extraction network based on Transform, mining a sensor time sequence dependency relationship from a time dimension, and analyzing multi-sensor association features from a space dimension; 3, introducing a particle filter algorithm to carry out optimal fusion on space-time dual-channel output, and quantifying the uncertainty of a data recovery result; and step 4, by comparing the actually measured data with the uncertainty confidence interval, abnormal detection and data recovery of sensor anomalies are realized. The invention belongs to the technical field of satellite sensor abnormal data recovery.

The invention discloses a simulation lunar dust excitation speed measurement system and method under a vacuum plume action condition, and relates to the technical field of aerospace engineering, and the system comprises a vacuum chamber and thruster system, an image collection assembly, a synchronous control collection system, a pulse laser light path, a high-speed spotlight, and a simulation lunar surface. In the invention, two optical systems of particle image velocity measurement and particle tracking velocity measurement are integrated in the same experiment, and the 532nm optical filter and the light trapping sheet are configured to realize compatibility of optical requirements, so that the problems that lunar dust particles with different scales and velocities cannot be measured synchronously and the optical systems interfere with each other in the prior art are solved; according to the method, velocity field data of micron-sized high-speed small particles and millimeter-sized low-speed large particles can be synchronously obtained and directly correspond to each other, the measurement integrity and reliability are improved in combination with simulation environment and synchronous control, and accurate data support is provided for spacecraft lunar dust prevention design.

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.