207 results about "Space debris" patented technology

The invention relates to the technical field of spacecraft orbit control and space safety, and discloses a space debris orbit prediction and avoidance method, which comprises the following steps that: a ground computing center screens candidate targets and generates enhanced data packets for uploading; the spaceborne computer combines real-time navigation data to carry out geometric feature screening, and an instant high-risk target list is generated; controlling the star sensor to execute optical observation, and resolving an optical measurement relative orbit state by utilizing displacement superposition and Kalman filtering; selecting a high confidence coefficient or a conservative probability threshold according to the optical measurement state acquisition condition, and generating an avoidance instruction when the collision probability exceeds the limit; and the propulsion system responds to the instruction to execute avoidance, verifies the effect by confirming the semi-major axis variation and executes track recovery by selecting an aircraft. According to the method, the ground-air collaborative screening and shift superposition enhancement technology is adopted, the problems that the low-precision ephemeris false alarm rate is high and the detection capability of a satellite-borne sensor on a dark and weak target is insufficient are solved, and the accuracy of evasion decision making and the execution reliability are improved.

The present disclosure is a device and methods for satellite trajectory control and collision avoidance. Embodiments of the disclosure are comprised of a smart satellite device performing a process three steps. First, sensors collect data about the satellite's physical landing environment, passing information to satellite's database and processors. Second, the processors manipulate the information with a deep reinforcement learning program to produce instructions. Third, the instructions steer the satellite body by manipulating the satellite's panels for optimal trajectory and collision avoidance. The purpose for the present disclosure is to help solve the space debris problem.

The invention provides a multi-station observation segmental arc association method and system for space debris. The method comprises the following steps: establishing an initial orbit of an unidentified segmental arc and an association information database; inputting an observation arc section; calculating an initial orbit of the observation arc section; traversing the initial track of the unidentified arc section and existing entries in the associated information database, sequentially judging whether the current observation arc section is associated with the initial track of the unidentified arc section and the existing entries in the associated information database or not until an associated object is found or all the existing entries are traversed, and if the associated object is found, judging whether the current observation arc section is associated with the existing entries in the associated information database. Modifying the initial track of the unidentified segmental arc and the content of corresponding items in the associated information database; if the associated object is not found, the information of the observed arc section is used as a new record to be inserted into the initial track of the unidentified arc section and the associated information database; the method is easy to implement, different association strategies are adopted for data of different time spans, association time consumption is reduced while association accuracy is guaranteed, and the method is suitable for association between single-station and multi-station observation arc section data.

The system and method for decelerating space objects moving at high velocities are provided. The system comprises at least one closed volume containing a combination of gas, liquid, solid particles, or a mixture thereof. The volume can be transported to the targeted location using a range of means, including a chemical gun, light gas gun, electromagnetic coil gun, superconducting quench gun, a rocket, or a combination thereof. The volume is strategically positioned on the trajectory of the moving object. Upon penetrating the walls of the volume, the object passes through it, experiencing deceleration.

Space debris capture systems include an outer assembly and an inner assembly. The Outer assembly includes an outer membrane and an inner membrane, where, in a deployed state, the outer membrane is separated from the inner membrane by a separation gap. The inner assembly is arranged within the inner membrane of the outer assembly. The inner assembly is formed of a plurality of sheets of material arranged at a plurality of orientations relative to the inner membrane and define a plurality of cells within the outer assembly when in the deployed state. A first group of the cells are defined within surfaces of the plurality of sheets of material and a second group of cells are defined between the surfaces of the inner assembly and an inner surface of the inner membrane.

The invention discloses a star-sky dim and small target detection method fusing coordinate attention and a convolutional feedforward neural network, and belongs to the technical field of dim target detection. The method comprises the steps of obtaining a starry sky image data set, outputting an enhanced feature map by applying a coordinate attention mechanism, constructing a network model fusing the coordinate attention mechanism and a ConvFFN module through the ConvFFN module of a convolutional feedforward neural network, training a verification model based on simulation data, and finally outputting the centroid position and bounding box information of a target. According to the method, features are extracted in the transverse and longitudinal directions through a coordinate attention mechanism, the feature expression ability of a weak and small target in a low signal-to-noise ratio environment is effectively enhanced, the detection precision is remarkably improved, the omission ratio is reduced, the calculation efficiency is improved by combining the lightweight structure of ConvFFN and deep convolution optimization, and the method is suitable for large-scale popularization and application. A reliable and efficient detection solution is provided for astronomical observation and space debris monitoring.

The embodiment of the invention provides a collision avoidance method and device, equipment and a storage medium, which can be applied to the technical field of space debris, and the method comprises the steps: determining a plurality of time intervals for a spacecraft to execute collision avoidance on at least one dangerous object within a set time period; for any time interval, determining at least one avoidance strategy in the time interval based on the initial orbit of the spacecraft and the operation orbit of the at least one dangerous object; the evasion strategy comprises a target pulse moment and a target evasion speed pulse applied to the spacecraft at the target pulse moment; predicting whether a dangerous object with a new collision risk exists or not after the at least one avoidance strategy is executed; and the avoidance strategy without the new collision risk is determined as the target avoidance strategy, so that accurate avoidance of at least one dangerous object by executing the avoidance strategy once is realized, and the collision avoidance calculation automation and intelligence are improved.

Container for collecting space debris. Space debris ingress through container's openable access port. A barrier deployment mechanism has a rotatable roll of flexible material on a frame assembly. A barrier translation mechanism translates the frame assembly towards or away from the access port to adjust the size of the compartment. When unwound, the roll forms a compartment for containing the space debris within the interior of the container, thus forming an obstruction between the compartment and access port.

The invention discloses a space target disintegration moment judgment method based on debris distance characteristics, solves the problem of how to improve the judgment accuracy of the space target disintegration moment, and belongs to the field of space target disintegration time inversion. Comprising the following steps: calculating the relative distance of each group of debris pairs in a target space debris set at any time through an Euclidean distance formula, and constructing an average relative distance equation; determining a time window according to prior information of a disintegration event, and iteratively solving a preliminary disintegration moment through a rough search method and a golden section method based on an average relative distance equation; through a density clustering algorithm, adaptively eliminating noise fragments at the initial disintegration moment, then reserving a main cluster fragment set, and constructing a corrected average relative distance equation of the main cluster fragment set; a local window is obtained with the initial disintegration moment as the center, the space target disintegration moment is iteratively solved through a rough search method and a golden section method on the basis of a corrected average relative distance equation, and the disintegration moment positioning efficiency is improved.

The invention relates to a method for calculating the degree of polarization of complex space debris, in particular to a method for calculating the degree of polarization of complex space debris based on a micro-facet pBRDF model, and solves the technical problem that an existing method for calculating the degree of polarization is poor in authenticity and accuracy. According to the complex space debris polarization degree calculation method based on the micro-facet pBRDF model provided by the invention, aiming at the characteristics of multi-shape and multi-material coupling of complex space debris, on the basis of the modeling thought of complex refractive index, specular reflection, diffuse reflection and volume scattering of different structural materials of the complex space debris; the albedo projection area is adopted as a reflection contribution coefficient calculated by the total light intensity of reflected light of the complex space debris, and a micro-surface element pBRDF model of different structures and different reflection contribution coefficients of the complex space debris is established, so that the model is closer to the multi-shape and multi-material characteristics of the real space debris; and the polarization degree of the complex space debris calculated based on the micro-facet pBRDF model is more accurate.

The invention provides a space station maintenance simulation system and method, and relates to the technical field of digital twinning, and the system comprises a space station environment simulation construction module which is used for constructing a simulation environment of a space station through a mixed reality technology according to the sun illumination, the atmospheric radiation in the morning and evening period and space debris orbit parameters; the model adaptability reconstruction module is used for performing three-dimensional model rendering on the white mold basic structure to obtain a space station model; the space station damage condition simulation module is used for constructing an impact damage three-dimensional image through the collected impact historical data between the space debris and the space station, and superposing the impact damage three-dimensional image to the space station model; the space station maintenance scheme selection module is used for recommending a mechanical arm maintenance scheme or an astronaut out-of-cabin maintenance scheme according to the impact damage three-dimensional image, the fault report condition and a real pre-judgment analysis strategy input by a user; simulation construction of a space station environment is realized on the basis of a virtual mixed reality technology, and the cost of extravehicular maintenance training of the space station is saved.

The invention belongs to the technical field of on-orbit spacecrafts, and relates to a large light deformable protective shield for an on-orbit spacecraft. The protective shield is simple and compact in structure, few in transmission path and high in overall reliability. The servo motor is adopted for control, the control precision is high, and remote control or remote sensing control can be achieved. The on-orbit spacecraft can be effectively protected by opening the protective shield, and impact of space debris on the on-orbit spacecraft is avoided or relieved.

The invention relates to the technical field of aerospace engineering, space robots and intelligent control, and particularly discloses a space debris intelligent capture and treatment method and system based on an SAC algorithm, and the method comprises the steps: obtaining the motion data of a space non-cooperative target and the system parameters of a space robot; constructing a hybrid filtering optimization framework based on the motion data and the system parameters; adjusting a hybrid weight and a noise forgetting factor in the hybrid filtering optimization framework by using an SAC algorithm to obtain self-adaptive estimation of a target attitude; a state space and an action space are defined based on a base of the space robot and a mechanical arm of the space robot respectively, and based on the state space and the action space, a mechanical arm movement path of the space robot is generated through an SAC algorithm in combination with a multi-target reward function; and the self-adaptive estimation of the target attitude is fed back to the motion path of the mechanical arm, dynamic cooperation of filtering and control is achieved, and stable capture of the space non-cooperative target is completed.

The invention discloses a high-storage-ratio folding method for a large conical film structure, and belongs to the technical field of space debris removal and satellite light shields. The problems that an existing plane deorbit sail is poor in attitude stability and low in deorbit efficiency are solved. Comprising the following steps: S1, drawing equidistant circumferential folding lines on a conical film structure, wherein the circumferential folding lines comprise a circumferential ridge line and a circumferential valley line; s2, circumferential folding is carried out, axial compression is carried out along the axis of the cone, and then circumferential contraction is carried out; and S3, folding from a cone to a cylinder. The three-dimensional conical film sail can be folded into the three-dimensional conical film sail which is high in posture stability, relatively high in unfolding reliability and capable of correcting the posture, and the folding efficiency is high.

The invention discloses a space debris target detection and positioning method based on dynamic programming and multi-hypothesis testing, and the method comprises the following steps: carrying out the preprocessing of an obtained space debris image, and obtaining a candidate point state set under different time frames; in the preprocessing, a denoising segmentation extraction channel and a lightweight CNN target extraction channel are divided according to a statistical local signal-to-noise ratio; obtaining a plurality of motion paths corresponding to each candidate point according to the candidate point state sets under different time frames; and for each candidate point, generating an initial hypothesis set according to the corresponding multiple motion paths, and processing the initial hypothesis set through Kalman filtering to obtain the position of the candidate point. According to the method, the detection precision of the dark and weak target is remarkably improved, the false alarm rate is reduced, the real-time performance of detection is improved, and the method is suitable for real-time monitoring scenes. The research result provides effective technical support for improving the operation safety of the spacecraft.

The invention provides a radar resource dynamic allocation method and system based on multi-task reinforcement learning, and relates to the technical field of multi-agent reinforcement learning multi-beam radar resource allocation, and the method comprises the steps: constructing a matrix through space debris orbit parameters, collision probability and radar detection historical data, and dividing the matrix into high, medium and low threat levels; then, a multi-beam radar system is configured, and different gain beams are distributed for the target according to the threat level; then, a distributed multi-agent reinforcement learning framework is adopted, an agent is deployed for each radar, the agents are combined with related data to generate a collaborative resource allocation strategy, and beam pointing and a scanning period are adjusted; and finally, integrating a threat-driven reward mechanism, calculating a reward weight according to the fragment collision probability, inclining resources to a high-collision-probability target, and realizing strategy adjustment, so that the detection and tracking efficiency of the radar on the space debris can be improved.

The application provides a variable configuration spacecraft system and method for non-cooperative target capture, and belongs to the field of deep space exploration. The method comprises the following steps: 1) a rigid combination flies to the vicinity of the target; 2) the combination is spun to be expanded into a ring belt; 3) the ring belt envelopes the target; and 4) the ring belt encloses the target. Large space debris and small celestial bodies both belong to non-cooperative targets, and cannot provide special interfaces for assisting capture. It is extremely difficult to develop general control mechanisms and methods to realize reliable capture and stable connection of various targets. In order to solve the problem, the application utilizes the inter-satellite connection to expand a plurality of controllable satellites into a ring-shaped multi-body tether satellite formation system, and carries out ring belt enveloping and then enclosing the tether to realize capture of a non-cooperative target with large scale, complex geometric shape, great uncertainty of motion state and inertia information according to a predetermined trajectory. The application has the advantages of high applicability, flexibility, high intelligence, reusability and small size.

The invention belongs to the technical field of aerospace science, and particularly relates to a lightweight target detection method suitable for a low-light space environment, and the method comprises the steps: 1, constructing a lightweight feature extraction backbone network based on MobileNetV4; 2, realizing multi-scale feature fine fusion in combination with TopkSPPF multi-branch pooling, and introducing a C2PSA attention module to improve the target focusing capability; 3, an EdgeFca module is embedded in the feature fusion stage, wavelet edge extraction and a frequency domain weighting mechanism are fused, and feature expression under low illumination is enhanced; and 4, designing a DNiN decoupling detection head and completing model training optimization. According to the method, the problems of low target detection precision, fuzzy edge, poor anti-noise capability, unsuitability for a satellite-borne edge platform and the like in a low-illumination space environment in the prior art are solved, and the method has a frequency domain modeling capability and an edge enhancement mechanism; and real-time identification and perception of non-cooperative targets such as spacecrafts and space debris can be realized under the condition of limited resources.

The invention relates to the technical field of space non-cooperative target capturing and racemization, in particular to a space non-cooperative target variable-stiffness racemization device based on a dielectric elastomer and application of the space non-cooperative target variable-stiffness racemization device, and the space non-cooperative target variable-stiffness racemization device is suitable for attitude control and racemization operation of space tumbling satellites, space debris and other non-cooperative targets. The supporting assembly comprises a middle shaft, a middle shaft servo motor and a fixing base, the middle shaft vertically penetrates through the fixing base, and the middle shaft servo motor is in transmission connection with the middle shaft and used for driving the middle shaft to rotate around the axis of the middle shaft or driving the middle shaft to pitch in the vertical direction so as to adjust the space posture of the arm rotation eliminating body. Compared with the prior art, the racemization device is high in racemization precision, good in operation flexibility, free of fuel consumption and suitable for long-term space tasks.

1. Name of the design product: Spacecraft (Future Design · Space Sweeper). 2. Use of the design product: For space debris cleaning, flying, and display. 3. Design points of the design product: In shape. 4. Picture or photo that best indicates the design points: Perspective view.

A system and method for decelerating space objects are provided. The system comprises at least one enclosure that is inflated by a medium, such as a gas, liquid, solid particles, or any mixture thereof. The enclosure can be transported to a predetermined orbital location in the path of the space object. Upon penetrating the enclosure membrane, the object passes through the medium and is thereby decelerated leading to decay of its orbit and / or disintegration thereof.

The invention discloses a low-orbit large spacecraft aluminum alloy framework autonomous disintegration system and a control method. The low-orbit large spacecraft aluminum alloy framework autonomous disintegration system comprises an aluminum alloy framework, an activation signal unit and a liquid metal triggering unit. The activation signal unit is used for receiving an external signal and sending an activation signal to the liquid metal triggering unit; the liquid metal triggering unit comprises liquid metal, a storage bin for sealing the liquid metal and a protective gate; the protective gate is arranged between the liquid metal and the aluminum alloy framework so as to isolate the liquid metal from the aluminum alloy framework; the protection gate receives the activation signal and releases the liquid metal, so that the liquid metal makes contact with the aluminum alloy framework and diffuses and permeates into the aluminum alloy framework along the grain boundary of the aluminum alloy, the structural integrity is destroyed, the spacecraft is decomposed, the orbit is lowered under the influence of residual atmosphere and the like, and finally the spacecraft enters the atmosphere to be completely burnt out. Compared with the prior art, the device can be controllably disassembled, and space junk is avoided.

The invention relates to the technical field of spaceflight, and particularly discloses a satellite combination structure and a satellite launching control method. The satellite combination structure comprises a first satellite and at least two second satellites; the second satellites are arranged on the same side of the first satellite side by side, the second satellites are connected with the first satellite through separation mechanisms, and every two adjacent second satellites are connected through a separation mechanism; the first satellite comprises a propelling system and a first force bearing structure, the propelling system comprises a main propeller and a fuel tank, the fuel tank is located in the first force bearing structure, and the main propeller is connected to the outer side of the bottom of the first force bearing structure. Therefore, complex structures such as an upper stage or a propulsion power cabin do not need to be additionally arranged, under the condition that the satellite function and the on-orbit deployment condition are guaranteed, the carrying capacity is effectively saved, new space junk is avoided, meanwhile, the satellite launching cost is saved, and the system construction cost is reduced.

The invention discloses a thermal unlocking structure for a spaceflight satellite. The thermal unlocking structure comprises a first structural part, a second structural part, a thermal unlocking assembly, a winding assembly and a restraint strap. After the restraint strap is fused by the hot knife, the knife holder retreats under the action of the spring, and the stop pin synchronously unlocks the limiting piece on the winding roller; the winding roller is driven by the pre-tightening torsion spring to rotate rapidly, and the fused restraint strap is completely rewound into the closed cavity in the shell through the guide wire. According to the design, it is ensured that restraint strap residues generated in the separation process are actively stored and restrained, and the risk that the restraint strap residues freely drift away to become space junk or interfere with a satellite body and an unfolding mechanism is fundamentally eliminated. The whole recovery process does not need extra energy, only depends on elastic potential energy pre-stored by the mechanism for driving, has the advantages of high reliability, low impact and no secondary pollution, and remarkably improves the safety and cleanliness of an on-orbit unlocking and releasing task of the spacecraft.

The invention provides a link state sensing dynamic service function chain deployment method in an F6G low earth orbit satellite optical network, and relates to the technical field of low earth orbit satellite optical network service function chain deployment. The method comprises the following steps: acquiring TLE data of space debris according to network topology information and network resource information of an F6G low earth orbit satellite optical network; preprocessing the TLE data of the space debris, and performing clustering processing by using a KMeans clustering algorithm; predicting the risk level of the risk fragment, and calculating the interruption duration of a risk link and a comprehensive risk value of the link; constructing a weighted satellite optical network topological structure based on the link logic distance; obtaining k optimal pre-deployment paths by using a K shortest path algorithm; constructing an optimization problem to carry out flow routing and VNF placement joint optimization; and if the constraint condition is met and no high-risk link causes business interruption, the service deployment is successful. According to the method, the acceptance rate and the resource utilization rate of the whole service of the system are effectively improved, and meanwhile, the average time delay of service deployment is reduced.

The invention provides a calculation method and device for space debris avoidance and electronic equipment, and the calculation method comprises the steps: calculating a reachable domain of a spacecraft based on a neural network of multi-source heterogeneous data modeling, and according to the state parameters of the at least one space debris, calculating an orbit range of the at least one space debris in the prediction time by adopting an orbit propagation model, calculating an overlapping region of a reachable region and the orbit range, and for the overlapping region, solving an evasion orbit of the spacecraft meeting constraint conditions based on a genetic algorithm. The autonomous collision avoidance capability of the spacecraft can be improved, the safe operation of the spacecraft in a complex space debris environment is ensured, and the fuel utilization rate of the spacecraft is improved.

The invention belongs to the technical field of spacecraft orbit design and operation management, and provides a satellite orbit life rapid calculation method, which combines a differential-integral method with Gaussian-Legendre integral, calculates the average change rate of orbit elements in an orbit period through the differential-integral method, and calculates the service life of the satellite orbit. The method combines the adaptive sampling characteristics of Gaussian-Legendre integral to realize large-step integral, greatly reduces the number of iterations and calculation amount, repeats the integral until the perigee height is less than the reentry height, the satellite falls down, and outputs the cumulative time of multiple integrals, namely the service life of the satellite orbit, thereby solving the technical problem that the calculation efficiency and precision cannot be considered at the same time. The method has important application value for satellite orbit maintenance, space debris management, reentry prediction and the like.

The present application relates to the technical field of space debris protection, and more particularly to a flying escort small satellite and a spacecraft operation system, comprising a first cabin body, a second cabin body, a third cabin body, a first protection assembly, a second protection assembly, a third protection assembly and a fourth protection assembly, which can perform both inspection and protection tasks, and improve the utilization degree of the flying escort small satellite. When space debris with a size of 1cm-10cm attacks, the second cabin body is extended out of the third accommodating cavity of the first cabin body, the third cabin body is extended out of the fifth accommodating cavity of the second cabin body, the first protection assembly is extended out of the first accommodating cavity, the second protection assembly is extended out of the second accommodating cavity, the third protection assembly is extended out of the fourth accommodating cavity, and the fourth protection assembly is extended out of the sixth accommodating cavity, thereby forming a multi-level defense barrier, playing a multi-level interception role on the space debris with a size of 1cm-10cm, and maximizing the fragmentation and kinetic energy dissipation of the space debris.

The application discloses a kind of non-cooperative target centroid intelligent positioning method and system fusing multi-dimensional physical constraint.The method is directed to sparse, noisy point cloud data, and constructs a comprehensive evaluation function containing four-dimensional information of geometry, orbital dynamics, time sequence continuity and surface physical properties;Using a two-stage solution framework of surrogate model and hybrid optimization, first, a lightweight neural network surrogate model is used with a differential evolution algorithm for global coarse search, and then switching to a high-fidelity orbit model combined with an adaptive particle swarm optimization algorithm for local fine search;Finally, the optimal state estimation and its covariance matrix, evaluation decomposition and other decision support information are output.The application improves the accuracy, robustness and computational efficiency of centroid positioning, and enhances the interpretability of the results, suitable for on-orbit servicing, space debris removal and other high-risk space missions.

The invention discloses a space debris protection structure for a spacecraft. The space debris dot matrix protection structure comprises a buffer plate, a dot matrix structure and a rear wall. The buffer plate can be made of metal materials such as aluminum alloy and titanium alloy, and the buffer plate is used for consuming kinetic energy of the space debris and breaking the incident space debris into debris cloud, so that a subsequent structure is easier to protect. The lattice structure is used for deflecting the kinetic energy direction of the fragment cloud so that the fragment cloud can impact more inclined plates, and therefore the kinetic energy of the fragment cloud is consumed. And the rear wall is used for protecting a small amount of fragments transmitted from the lattice structure. The space debris protection structure provided by the invention can improve the defense capacity of a spacecraft to space debris, has certain bearing capacity, and compared with a traditional Whipple structure, the protection performance of the space debris protection structure is improved by 37% and the space is reduced by 25%-50% under the same surface density, and the space debris protection structure has the characteristics of saving the installation space, being light in weight, being capable of bearing and having good heat insulation and protection effects.