183results about "Cosmonautic vehicle tracking" patented technology

In a space traffic management system (500), space traffic management devices (100) individually mounted in a plurality of mega-constellation business devices and in a debris removal business device (45) are connected to each other via a communication line (200). The debris removal device (45) performs Active Debris Removal (ADR) against debris formed by a satellite managed by a first mega-constellation business operator. The debris removal device (45) acquires real-time high-accuracy orbital information of a satellite group of a second mega-constellation business operator in a timeframe in which a debris removal satellite, during orbital descent, passes through an orbital altitude region where the satellite group of the second mega-constellation flies, the debris removal satellite passing through the satellite group while ensuring flight safety.

Various embodiments comprise a multi-agent spacecraft system for rendezvous and proximity operations. The spacecraft system comprises a chief spacecraft and a deputy spacecraft. The chief spacecraft maneuvers to the orbit of a Resident Space Object (RSO) and deploys a deputy spacecraft to rendezvous with the RSO. The deputy spacecraft maneuvers to the RSO and senses the RSO during the rendezvous and proximity operations maneuver. The deputy spacecraft generates a Three-Dimensional (3D) surface reconstruction of the RSO based on the sensing. The deputy spacecraft determines a roll rate and a relative trajectory of the RSO based on the sensing. The deputy spacecraft returns to the chief spacecraft.

A satellite constellation forming system (100) forms a satellite constellation (20). The satellite constellation (20) is composed of a satellite group (300). In the satellite constellation (20), the satellite group (300) provides a service cooperatively. The satellite constellation (20) has a plurality of orbital planes in which a plurality of satellites (30) fly at the same orbital altitude in each orbital plane (21). A satellite constellation forming unit (110) forms the satellite constellation (20) in which orbital altitudes of the plurality of orbital planes (21) are mutually different.

The disclosed technology is generally directed to a method for simulating light pollution by a satellite system including at least one satellite. In one example of the technology, the method may include receiving orientation and solar panel parameters associated with the satellite, a position of the Earth with respect to the Sun, importing a database including multiple point of interests (POIs) on the Earth, and determining a surface roughness parameter of the satellite using a Gaussian distribution model. Based on the orientation and solar panel parameters, the position of the Earth, and the surface roughness parameter: simulating the light pollution by the satellite to iteratively determine the light pollution at the POIs. The method may include storing the light pollution determined at the POIs in a buffer and dynamically rendering a user interface to display a visualization representing the light pollution by the satellite at the POIs.

The application discloses a spin-nutation collaborative inhibition method and system for a space tumbling target, and belongs to the technical field of aerospace. The method comprises the following steps: acquiring pose data of a space robot, and controlling the space robot to approach the space tumbling target according to the acquired pose data of the space robot; after the space robot approaches the space tumbling target, the space robot measures the space tumbling target, and acquires real-time state data of the space tumbling target by measuring the space tumbling target; a target stress position prediction model of spin-nutation collaborative inhibition is used to simulate and predict based on the acquired real-time state data of the space tumbling target, and a despinning condition required by a despinning tool in the space robot is determined; the space robot adjusts the pose according to the despinning condition, and controls the despinning tool to contact the space tumbling target, and implements contact despinning. The application can efficiently inhibit the nutation of the target in the initial stage of the despinning task, and effectively reduces the difficulty of the subsequent despinning task.

In a space traffic management system (500), space traffic management devices (100) individually mounted in a plurality of mega-constellation business devices and in a debris removal business device (45) are connected to each other via a communication line (200). The debris removal device (45) performs Active Debris Removal (ADR) against debris formed by a satellite managed by a first mega-constellation business operator. The debris removal device (45) acquires real-time high-accuracy orbital information of a satellite group of a second mega-constellation business operator in a timeframe in which a debris removal satellite, during orbital descent, passes through an orbital altitude region where the satellite group of the second mega-constellation flies, the debris removal satellite passing through the satellite group while ensuring flight safety.

A satellite tracking system and method, a three-point peaking technique is used to determine the direction of a satellite 50 which includes a signal source. A Kalman filter is used to minimize the effects of noise in the received signal during the three-point peaking operation. This determination may be made at any time. Once the position of a satellite 50 has been determined twice over a time interval, its future position may be estimated by an adaptive continuous step track technique, using a Kalman filter, which assumes that the satellite moves uniformly with time as viewed from the antenna.

A space traffic management device (100) mounted in a mega-constellation business device (41) specifies one to a plurality of representative satellites from a mega-constellation satellite group flying on orbits having the same nominal orbital altitude, and has quasi-real-time high-accuracy orbital information of the representative satellite and orbital information relative values of constituent satellites other than the representative satellite, relative to the representative satellite. The space traffic management device (100) shares the quasi-real-time high-accuracy orbital information of the representative satellite and the orbital information relative values of the constituent satellites relative to the representative satellite, with the space traffic management devices (100) mounted in a plurality of business devices.

A system for earth monitoring includes a constellation of satellites (310) and a dispatch computing module (302) located on earth. The dispatch computing module (302) can be configured to implement a satellite management method that receives a post-launch request for a location to be imaged; re-computes an existing schedule of satellites to provide an updated schedule that includes imaging the location; and provides the updated schedule to one or more of the satellites for transmission.

This invention provides a spatial target tracking method based on an optical search component, comprising the following steps: Step 1, rotating the field of view of the optical search component across its entire range to locate the target object; Step 2, transmitting the scene in the field of view in real time from the optical search component, processing the transmitted scene in the field of view, and determining whether the optical search component has tracked the target object; Step 3, establishing a predicted trajectory function for the target object based on its actual motion trajectory to obtain the predicted motion trajectory of the target object; Step 4, rotating the optical search component according to the predicted motion trajectory of the target object to adjust its field of view angle.

In a system where multiple satellites autonomously take turns acting as servers, the servers will be rotated according to instructions from a ground station. [Solution] The server, which is a satellite 200 that has obtained server rights, receives a priority notification when a ground station 300 transmits the priority notification. If the server receives the priority notification, it selects the satellite 200 specified in the priority notification as the server for the next time interval. If the server does not receive the priority notification, it selects one of the multiple satellites as the server for the next time interval based on the status information of the server and the status information of each client.

Examples of the present disclosure include a method for approximating a non-carth imaging camera model for a non-earth sensor includes obtaining a physical camera model, receiving a target range value, determining a plurality of space-based coordinates of a plurality of points about a line of sight of the sensor in proximity to a distance corresponding to the received target range value, for each of the plurality of locations, determining line and sample coordinates on an image thereof via the physical camera model based on the space-based coordinates of the plurality of points, and based on the determined line and sample coordinates for each of the plurality of points, fitting an approximate camera model thereto.

In some implementations, a device may obtain historical orbital data indicative of orbital movement of a set of satellites over a period of time, the set of satellites comprising at least a subset of satellites of the satellite constellation. The device may, for each satellite in the set of satellites, estimate a set of orbital parameter values by fitting an orbital model to the historical data of the respective satellite, wherein at least one orbital parameter value of the set of orbital values is common across a group of satellites within the set of satellites. The device may send assistance data to at least one mobile device, the assistance data indicative of the respective set of orbital parameter values for all satellites of the set of satellites.

Angular velocity control device for a spacecraft and corresponding spacecraft. The invention relates to a device (1) for controlling the angular velocity of a decommissioned spacecraft, comprising: - a stator (3) and a rotor (4) movable about an axis (A21) of rotation relative to the stator. The stator (3) comprises an electrically conductive and non-ferromagnetic body (6), while the rotor (4) comprises a magnetized system (7) configured to induce, in the stator (3), braking eddy currents of a relative motion of the rotor (4) with respect to the stator (3); - a levitation magnet (11) intended to cooperate with a magnetic field generated by an external source to induce magnetic levitation of the rotor (4) with respect to the stator (3). The device (1) is made of one or more non-ferromagnetic materials in a zone (11ZI) influenced by the magnetic field generated by the levitation magnet. SAMPLE FIGURE: [Fig.4]

To enhance versatility in analysis of acceleration and deceleration of a space object.SOLUTION: The acceleration / deceleration analyzing device 700 is included in a space situation monitoring business device that monitors a space object with a monitoring device, analyzes the orbit of the space object 60 to determine whether the space object 60 is accelerating or decelerating, and, when the monitoring device finds a space object A whose orbit information is unknown and is composed of elements of the orbit 6 based on the original t0 and Kepler's law, sets the following parameters as a parameter set 0 indicating an estimated orbit information initial value that is an estimated value of the orbit information of the space object A in the original t0: Monitoring information of a monitoring device about a space object A is acquired three times or more to derive orbit 6 elements in an original period t0. The acceleration and deceleration analyzer 700 also derives the acceleration of the space object A in the out-of-plane direction at the ascending node and the acceleration of the space object A in the out-of-plane direction at the descending node from the difference between the orbit inclination derived from the monitoring information about the space object A and the orbit inclination of the parameter set 0.SELECTED DRAWING: Figure 1

A monitoring system (1) for an orbiting space object (5) comprising a ground-based radar system (2) (7) including an antenna (6) with a controllable pointing direction for moving a space observation window (11) within a field of view (12), the pointing direction having a predetermined value (Zref) associated with the space object (5), the monitoring system (1) comprising ground-based optical sensors (3) (7) and a control unit (4) configured to determine a first orbital restitution (O1) from first positions (P1) of the space object (5) measured by the optical sensors (3), and, if the first orbital restitution (O1) crosses the field of view (12) of the radar system (2) outside the space observation window (11) associated with the predetermined value (Zref) of the pointing direction, to determine a command for the pointing direction of the antenna (6). Abstract figure: Figure 1

A satellite (140) for operation in an orbit around the Earth includes an ADCS (131, Fig. 1) configured to mechanically steer the satellite in an azimuth direction to extend a dwell time (1105) during which a target is visible from the satellite when the satellite is orbiting above the target of choice. A processor at a ground station can be configured to process raw SAR data from any of the satellites described herein. The raw SAR data can be processed in a variety of ways to provide image information, including but not limited to forming multi-view images, compiling video sequences, and color coding images.

In one embodiment, a user device (UE) is provided, comprising: a transceiver that receives system information during operation; and a circuit that, during operation, determines a first information element from the system information that identifies the setting parameters of a first satellite; and determines a second information element from the system information and, with reference to the setting parameters of the first satellite, that identifies the setting parameters of a second satellite different from those of the first satellite. A corresponding base station and a method including steps performed between the UE and the base station are also provided.

Systems, devices, methods, and computer-readable media for space vehicle security are provided. A space vehicle can include a navigational controller configured to determine a mode of the space vehicle, a power supply configured to provide power to components of the space vehicle, a transceiver configured to determine whether a communications channel to a ground station is open, and an intrusion detection monitor configured to monitor for non-nominal outputs from components of the space vehicle, the components that are monitored are variable based on a power level of the power supply, a connectivity state of the transceiver, and the mode.

The application provides an autonomous star observation method for super low orbit satellites, comprising: setting a whole satellite mode word; judging whether an autonomous working condition trigger enable flag is enabled according to the whole satellite mode word; if yes, judging an orbit height decision working condition, and starting a star observation task according to the working condition; if no, setting a working condition through a ground instruction, and starting a star observation task according to the working condition. Meanwhile, an autonomous star observation system for super low orbit satellites is provided. The application can adapt to star observation task requirements under super low orbit height conditions, can overcome large aerodynamic interference moments under super low orbit conditions and large rotation moments caused by a turntable motor, can autonomously switch super low orbit star observation working conditions according to orbit heights, can reduce errors caused by ground orbit prediction, and can retain the ability of ground intervention whole satellite tasks, and can switch whole satellite autonomous observation and ground forced observation at any time.

A method for correcting an orbit error of common-view time comparison of a space station, comprising five parts, i.e., determination of moments when two stations can simultaneously view the space station, calculation of projection coefficients of the orbit error, calculation of a three-dimensional orbit error vector set, calculation of a set of projection differences of the orbit error in line-of-sight directions of the two stations, and threshold decision. By means of calculation of the five parts, the maximum influence quantity of the orbit error on a common-view time comparison result is obtained, observation moments with the maximum influence quantity exceeding a decision threshold are eliminated, and observation points with the influence of the orbit error on the common-view time comparison less than the decision threshold are reserved.

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 embodiment of the invention provides a space target maneuvering detection method and system based on an information energy momentum tensor, and relates to the technical field of spaceflight detection. Comprising the steps of obtaining orbital elements of an observation satellite and a space target; according to the number of the orbits, the number of the wide-area invariant opposite orbits is obtained; obtaining a space target relative position after dimension normalization under an observation satellite LVLH coordinate system according to the wide-area invariant relative orbital element; according to the relative position and the number of the relative orbits, the relative motion acting amount and the relative speed are obtained; carrying out filtering processing on the relative motion state of the observation satellite according to a mean filtering algorithm; acquiring a corresponding energy momentum tensor according to the filtered relative motion action quantity, relative position and relative speed; and obtaining a time-varying curve of the energy momentum tensor, and carrying out maneuvering detection on the space target. According to the method, the detection capability on low thrust and weak signal maneuver can be remarkably enhanced, and the limitation that a traditional method is high in dependence on model precision and insensitive to micro maneuver is overcome.

To achieve information on a mission device for executing mission other than communication via annular communication network in real time at low cost.SOLUTION: A hybrid constellation 20B formed in a Low Earth Orbit includes a communication constellation in which a plurality of satellites having a communication device communicating with precedent and following satellites in a travel direction on the same orbital plane form annular communication network 702, and a mission satellite 30b which has a communication device communicating with the precedent and following satellites, and a mission device for executing various missions. The mission satellite 30b flies between the plurality of satellites forming the communication constellation, the hybrid constellation 20b is formed by reconstruction of the annular communication network 702 using a plurality of satellites forming the communication constellation forming the annular communication network 702 and the mission satellite 30b.SELECTED DRAWING: Figure 17

Various examples are provided related to a space-based imaging approach for characterizing objects in space Earth or other planetary body. In one example, a method for characterizing objects in space includes illuminating an object in space about Earth or other planetary body with a narrow band, continuous wave (CW) radio beam transmitted by a transmitting telescope; receiving return signals from the object by receiving telescopes in an array, at least one receiving telescope in orbit; and generating a high resolution image of the object from the received return signals utilizing a near-field correction. In another example, a space-based imaging system includes an array of telescopes including a transmitting telescope that can illuminate an object in space; a plurality of receiving telescopes that receive return signals reflected by the object; and processing circuitry that generates a high resolution image of the object from the received return signals.

The ever increasing number of orbiting bodies in low Earth orbit has made it infeasible to calculate potential conjunctions between orbiting bodies more than a few days in advance, even with the aid of supercomputers. Disclosed embodiments utilize machine learning to predict potential conjunctions between orbiting bodies faster than state-of-the art systems by orders of magnitude. This enables potential conjunctions to be identified well in advance (e.g., 30 days or more), so that they may be prioritized (e.g., for fine calculations), visualized, and remediated (e.g., via control of the impacted satellites).

To enhance versatility in analysis of acceleration / deceleration of a space object.SOLUTION: An acceleration / deceleration analysis device (700) is provided on a space condition monitoring business device for monitoring a space object with a monitoring device, analyzes an orbit of a space object (60) and analyzes presence / absence of acceleration / deceleration of the space object (60), and acquires monitoring information on the monitoring device for a space object A, in which the orbit information composed of the original period t0 and the orbit 6 elements based on Kepler's law is unknown, three times or more, as a parameter set 0 indicating an estimated orbit information initial value that is an estimated value of orbit information of the space object A at the original period t0, when the monitoring device discovers the space object A, and derives orbit six factors at an original period t0.SELECTED DRAWING: Figure 1