125 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.

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 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 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 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 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 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 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 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.

The invention discloses a high-orbit multi-source point target cross-domain fusion method based on modal characteristic distillation, and relates to the technical field of modern aerospace monitoring. The method comprises the following steps: extracting invariance features related to target essential attributes from visible light, infrared and radar data through a teacher network, and fusing the invariance features to generate teacher knowledge vectors; and taking the teacher knowledge vector as an optimization target of a student network, and performing iterative optimization by adopting a composite loss function including knowledge distillation loss, physical attribute loss, modal alignment loss and task loss, thereby realizing accurate identification of a point target. Through the feature extraction and knowledge distillation mechanism guided by physical constraints, the problem of feature heterogeneity among multi-modal data is effectively solved, the recognition accuracy and tracking stability of point targets such as space debris in a complex space environment are improved, meanwhile, the model is lightweight and efficient, and the on-orbit real-time processing requirement is met.

The invention relates to a satellite system (100) for monitoring and detecting space objects (102) in near-celestial space, in particular for monitoring and detecting space debris, preferably for determining at least one property of the space objects (102), comprising satellites (108, 110, 112, 114, 160) arranged on at least one first principal orbit (104), a torus-shaped monitoring area (116) extending around the celestial body, wherein each of the satellites (108, 110, 112, 114, 160) has an observation device (164) arranged and configured to generate images of space objects (102) within an observation field (108b, 110b, 112b, 114b) and to image space objects (102) within the torus-shaped monitoring area (116), at least one Control device (166) which is arranged and designed to align the observation fields (108b, 110b, 112b, 114b) in such a manner,that images of the space objects (102) located in the torus-shaped monitoring area (116) are generated.

The invention relates to a multi-source electromagnetic matrix aircraft take-off and landing system and a waste heat utilization and building integration method thereof, and belongs to the technical field of aircraft take-off and landing. The system comprises at least three electromagnetic emission units which are arranged in a space array, and each electromagnetic emission unit comprises an independent controllable excitation coil; the central controller is used for independently adjusting current parameters of each electromagnetic emission unit; the at least three electromagnetic emission units act together to form at least four independent and controllable electromagnetic stress action points in a predetermined space area. The system comprises a building facade integration system, a waste heat recovery system and a military application system (space debris cleaning and aerospace vehicle launching). The system has the multiple advantages of high control dimension, high magnetic field intensity, low cost, mobility, building embeddability, comprehensive energy utilization, military and civilian dual-purpose, military defense value and the like, and can be used as a general take-off and landing infrastructure in the fields of urban air traffic, military aviation, spaceflight launching, military defense and the like.

The application discloses a spacecraft autonomous avoidance of space debris based on a deep deterministic policy gradient algorithm, classifies a scene, constructs different training environments according to different types of orbit threats, constructs a constrained Markov decision process, makes avoidance decisions while satisfying multiple constraints of the spacecraft, in an offline training stage, makes the intelligent agent interact with different types of environments continuously, and learns and trains corresponding neural networks, in an online application stage, first, threat identification is performed, corresponding trained neural networks are extracted, avoidance actions are quickly generated, the spacecraft can learn to make corresponding optimal decisions in different scenes and environment states, and the avoidance of space debris is realized online. The technology can enable the spacecraft to face space debris threats, autonomously formulate avoidance strategies, autonomously complete avoidance actions, autonomously grasp the timing and strength of space debris avoidance, and effectively respond to the space debris threats at the minimum cost (minimize interference with task execution).

Provided is a method, system, and computer program product for determining a degree of impact caused by space debris. A processor may receive data from a data stream of one or more satellites, the data including observational data of one or more objects in orbit. The processor may determine, based on an analysis of the observational data, that the one or more objects are space debris. The processor may identify, based on the analyzed observational data, ownership of the one or more objects that are determined to be space debris. The processor may determine, using one or more smart contracts, a degree of impact caused by the one or more objects that are space debris, where the degree of impact is attributed to an identified owner of the one or more objects.

The invention relates to a satellite system (100) for monitoring and detecting space objects (102) in the area of space close to a celestial body, in particular for monitoring and detecting space debris, preferably for determining at least one object property of the space objects (102), comprising satellites (108, 110, 112, 114, 160) located on at least one first main orbit (104). Each of the satellites (108, 110, 112, 114, 160) has an observation device (164) which is mounted and designed so as to generate images of space objects (102) within an observation field (108b, 110b, 112b, 114b), and the at least one first main orbit (104) has an orbit height (178, 182) between 200 km and 800 km such that space objects (102) to be observed are preferably located at the same orbit height or higher orbit heights (178, 182) in order to align the observation field (108b, 110b, 112b, 114b) away from the celestial body (106).

The present application relates to the technical field of aerospace artificial intelligence, and provides a space debris collision risk probability prediction method based on deep learning, which takes ESA Kelvins real rendezvous data as a benchmark, and uses Kessler to generate a synthetic CDM sequence to expand the training sample to improve the coverage of high risk and complex situations; at the same time, a plurality of update records and their associated related events are expressed as a dynamic heterogeneous graph, a dynamic graph neural network is used to learn the law of risk evolution over time and the similarity between events can be optionally used for knowledge transfer, so that the continuous regression prediction result of the final collision risk probability of the rendezvous event can still be stably output under the condition of missing and noise, and the prediction is quickly refreshed when new CDM arrives through the local graph incremental update mechanism, which significantly improves the accuracy, robustness and online availability of risk prediction.

The invention discloses a small space debris-oriented passive capturing and disposing spacecraft, cubesat and method, solves the problems that in the prior art, small space debris cannot be captured, and an existing capturing device performs active capturing, and has the beneficial effects that the space debris is passively captured and thoroughly eliminated. According to the specific scheme, the passive grabbing and handling spacecraft for the small space debris comprises a chip satellite and a capturing assembly, a plurality of solar sails are connected to the peripheral side of the chip satellite, the chip satellite controls the bending state of the solar sails so as to control flight of the spacecraft, and the capturing assembly comprises a folding storage wall plate, a triggering platform and a bistable capturing arm; the folding storage wall plate is connected to the peripheral side of the chip satellite through the hinge, the hinge is a bent hinge and has elasticity, the triggering platform is connected with the chip satellite, and after the bistable capturing arm passes through the triggering platform, the bistable capturing arm is in a straight state when a spacecraft is in a non-working state.

This invention discloses a method for planning a spacetime flight path for launch vehicle debris avoidance, relating to the field of launch vehicle space debris avoidance technology. The method includes: establishing a directed graph model of the flight path based on the flight window at each altitude level; determining constraints based on the directed graph model, including graph search constraints and flight path constraints; combining the constraints with a debris avoidance trajectory planning problem model to establish a flight path planning model; the debris avoidance trajectory planning problem model is a model concerning rocket dynamics constraints, orbital insertion constraints, and debris avoidance constraints; and solving the flight path planning model to obtain the optimal flight trajectory. This invention significantly reduces launch vehicle fuel consumption through global optimization of the flight path; simultaneously, the method can be solved in real time, improving the timeliness and reliability of launch vehicle launches, and has certain engineering application potential.

The invention discloses a space debris detection method based on fluctuating phase parametric modeling. According to the invention, a fluctuating phase parameterized model is established based on space debris motion characteristics, a distance history is accurately extracted through generalized Rayden Fourier transform, then a three-dimensional fluctuating phase parameter space is constructed, long-time coherent accumulation is carried out on radar echoes in combination with Doppler phases corresponding to the distance history, and the radar echoes are obtained. And finally high-precision estimation of the fluctuating phase parameters of the space debris is realized. The invention aims to provide a high-precision space debris fluctuation phase parameter estimation method, accumulation loss caused by periodic phase fluctuation can be effectively compensated, higher coherent accumulation gain is obtained, and the detection probability of space debris under the condition of low signal-to-noise ratio is improved.

The invention discloses a navigation mode transition method, an electronic device and a medium, in a working process of a first navigation mode, geometric angles among a mother satellite, a service spacecraft and space debris are monitored, and when the geometric angles meet a mode switching requirement, the position of a target orbit under a J2000 coordinate system is obtained through double-satellite geometric estimation; and taking the position as an initial position calculated in the second navigation mode, and switching to the second navigation mode. Before the navigation mode is switched, the target orbit position is determined through double-satellite geometric estimation, the single-satellite navigation positioning precision can be rapidly improved, a good initial value is provided for double-satellite angle measurement filtering, and the convergence process is accelerated.

The application relates to the technical field of space debris capture, in particular to a device for recovering debris in a hypervelocity impact experiment, which comprises a target chamber and a capture structure, the target chamber is internally formed with a target cavity for placing a simulation target; the capture structure is arranged on the inner wall of the target cavity and comprises a first capture layer and a second capture layer, the second capture layer covers the inner wall of the target cavity, the first capture layer covers one side of the second capture layer away from the inner wall of the target cavity, the density of the first capture layer is smaller than that of the second capture layer, and the first capture layer is a transparent structure. Small-sized and low-energy micro debris are captured by the first capture layer, large-sized and high-energy debris pass through the first capture layer and are captured by the second capture layer, the number of micro debris entering the second capture layer is reduced, thereby facilitating the positioning and extraction of large debris in the second capture layer based on X-ray and sectioning methods, the debris recovery rate is improved, and the difficulty of extracting micro debris is reduced due to the transparent structure of the first capture layer.

The present invention provides a single-layer aluminum alloy plate for spacecraft that can reduce the number of ejectors generated when colliding with space debris, and a bumper for spacecraft made from this aluminum alloy plate. [Solution] The single-layer aluminum alloy sheet for spacecraft is used for the bumper of the spacecraft. The aluminum alloy sheet for spacecraft has a chemical composition consisting of Si: 1.0% to 8.0% by mass, with the remainder being Al and unavoidable impurities. The bumper for spacecraft, made of the single-layer aluminum alloy sheet for spacecraft, is attached to the outside of the spacecraft body.

The invention relates to the technical field of space debris orbit forecasting, in particular to a space debris orbit forecasting correction method based on machine learning and triangular characteristic decomposition. The azimuth angle and the pitch angle are converted into six-dimensional feature vectors through triangular feature decomposition, and a real-time mapping model of orbit forecast deviation is established in combination with a neural network. The two bottleneck problems on which traditional image tracking depends are solved: 1) in a strong laser interference scene, tracking failure caused by background light saturation of an optical camera is avoided, and real-time image feedback is replaced by deviation prediction driven by historical echo data; 2) when space debris in an earth shadow area is invisible, constructing a training set by using deviation data captured at the moment of successful laser ranging, and realizing autonomous correction of forecast errors under a non-illumination condition; the method gets rid of dependence on space debris optical visibility, realizes all-day and all-arc telescope pointing deviation correction, and improves the reliability and continuity of space debris orbit prediction under complex working conditions.

This invention belongs to the field of low-Earth orbit remote sensing satellite technology and provides a method and system for inter-satellite phase design based on sensor field-of-view matching. The method includes: observation and positioning to obtain the characterization of inter-satellite phase design input variables; calculating the upper and lower limits of the inter-satellite phase difference based on the characterization of the inter-satellite phase design input variables; determining the range of inter-satellite phase difference values; and outputting the phase difference. This invention solves the problem of field-of-view matching during satellite relay, improving the accuracy of relays. It further considers satellite sensor field-of-view matching when satellites have a cooperative relationship, making the problem more targeted to specific scenarios. It is applicable to applications such as Earth remote sensing observation and space debris observation, and is suitable for various sensors, including optical and microwave sensors. It can provide direct guidance for the design of various specific constellation schemes with inter-satellite relay cooperation.