33 results about "Orbital maneuver" patented technology

The invention relates to a spacecraft maneuver detection method based on TLE data and kernel density estimation. The invention relates to the technical field of spacecraft maneuver detection. The method comprises the following steps: acquiring and inputting satellite historical TLE data; gaussian filtering denoising is carried out on the input data; based on the de-noised data, propagation forecasting is carried out through an SGP4 model; constructing a joint residual sequence, and carrying out error distribution modeling and sample scoring; determining a maneuvering threshold value, and judging to obtain abnormal data; and aggregating abnormal data and carrying out maneuvering identification. The method can still realize accurate and stable track maneuver automatic detection under the constraint conditions of unstable TLE data, complex error distribution, no external auxiliary data and the like. The method has good adaptability, robustness and interpretability, and is a track anomaly detection technical scheme which is complete in structure, rigorous in logic and capable of achieving engineering landing.

Method, electronic device, system, and computer-readable medium embodiments are disclosed. Some embodiments include a signal processing workflow incorporating a graphical user interface for displaying orbital information for satellites and other spacecraft. In some embodiments, a generative adversarial network (GAN) is employed for evaluating satellite orbital positions, for predicting future orbital movements, for detecting orbital maneuvers of a satellite, and for analyzing such maneuvers for potential nefarious intent.

The invention discloses a method and a system for planning an orbital task of a maneuvering satellite, and belongs to the field of spaceflight simulation analysis. The method comprises the following steps: creating a planning task and setting basic parameters of a maneuvering satellite; setting basic parameters of a target satellite and a general control center position point; maneuvering analysis is carried out with maneuvering satellite parameters, target satellite parameters and the position point of the master control center as constraints, and a task time window meeting the constraints is obtained; after the task time window is determined, an orbital maneuvering multi-strategy algorithm is called for calculation to obtain a maneuvering mode and a maneuvering parameter list of the maneuvering satellite arriving at the task point; and finally, entering a planning event processing module, sorting events according to time, judging whether impact events exist or not, and outputting a planning task success / failure state. The general control center evaluates the fuel remaining amount in real time based on a fuel formula and supports a user to dynamically formulate and issue an orbital transfer scheme, and integrated planning of maneuvering satellite task window calculation, maneuvering strategy selection and safety check is achieved.

The invention provides a low-orbit constellation configuration maintenance method using perturbation force and adjacent position relative station deviation, and belongs to the technical field of low-orbit constellation configuration maintenance and control. The method comprises the following steps: acquiring orbit parameters of each satellite and adjacent satellites thereof in a low-orbit satellite constellation in a space, and calculating a relative drift distance and an allowable drift range of each satellite; on the basis of a low-orbit satellite constellation dynamic model and the relative drift distance, calculating theoretical retention time of each satellite in an allowable drift range, and further determining a constellation control sequence and an optimal control moment of each satellite; and when the satellite reaches the optimal control moment, determining the control quantity of the satellite orbit parameters and the propulsion system parameters of the satellite during orbit maneuver so as to maintain the constellation configuration. According to the method, the relative standing position relation of the adjacent satellites serves as control input, autonomous stable maintenance of the constellation relative configuration is achieved by controlling the orbit height and the inclination angle of the satellites, and unnecessary maintenance control and fuel consumption are avoided.

The embodiment of the present application discloses a kind of super low orbit satellite orbit height autonomous maintenance method.In a specific embodiment, the method comprises: the position and speed of satellite-borne GNSS receiver output as input, calculate the actual value of satellite's intersection period when satellite is in ascending node or descending node, according to the relationship between intersection period actual value and intersection period target value, execute orbit maneuver, to maintain orbit height.A kind of super low orbit satellite orbit height autonomous maintenance method of the present application, with the position and speed of satellite-borne GNSS receiver output as input, is realized by combination of electric propulsion and chemical propulsion, can comprehensively utilize the advantages of less electric propulsion working medium consumption and large chemical propulsion thrust, to realize orbit height on-orbit autonomous orbit maintenance.

A solar sail includes a bus and a plurality of separate movable vanes coupled to the bus. Each movable vane includes a reflective surface for generating solar radiation pressure and propel the solar sail in space. Each vane may be movable relative to the bus in a fully deployed configuration such that an amount of thrust generated by solar radiation pressure on each vane is controllable.

The invention provides an angle-measurement-only multi-pulse self-adaptive closed-loop rendezvous guidance method based on second-order cone programming, and belongs to the relative navigation and rendezvous guidance technology of spacecrafts. According to the method, firstly, an angle-only-measurement relative navigation dynamical model and a measurement model based on a CW equation are constructed under LVLH coordinates, system state observability analysis is carried out, the orbit maneuvering condition enabling the state to be observable is obtained, and the condition is used for constructing an optimal guidance index. Secondly, analyzing and modeling the rendezvous guidance error source, and deducing to obtain a relative orbit guidance error covariance model through error covariance analysis; thirdly, on the basis of a second-order cone programming method, guidance fuel consumption and system state observability conditions are fused, and a multi-target weighted cost function of optimal rendezvous guidance is jointly constructed; meanwhile, based on the established guidance error covariance model, a cost function weight index system with a self-adaptive adjustment function is designed, and real-time self-adaptive adjustment of a guidance strategy along with change of system performance is achieved. The method solves the problems of high fuel consumption, poor flexibility and the like caused by separate design of the relative navigation and the rendezvous guidance model in the traditional only-sight-line measurement rendezvous guidance method, and has the ability of adaptively adjusting and optimizing the rendezvous fuel consumption and the relative navigation system observability.

The invention provides a satellite launching / recycling system capable of moving back and forth freely and being loaded on orbit, and relates to the technical field of spaceflight. The system comprises a spacecraft which can be in on-orbit butt joint with a preset space station; the spacecraft comprises a three-section cabin body which comprises a front-section cabin body, a middle-section cabin body and a rear-section cabin body; a head fairing capable of being opened and closed is arranged at the end part of the front section cabin; a butt joint opening is formed in the head fairing. The top facing part of the middle cabin body is configured to be a cabin door capable of being opened and closed; an expandable folding solar wing and at least one satellite parking device are arranged on the inner side of the cabin door; a satellite parking point is formed after the cabin door is opened; a loading space used for loading a satellite to be launched and / or a recycled satellite, a mechanical arm device capable of grabbing the satellite and a satellite on-orbit launching device are arranged in the middle cabin body. The spacecraft is used as a carrier for satellite launching / recycling and / or a carrier for implementing orbital maneuver and returning to the ground in orbit. Satellite launching and recycling can be achieved with low cost and high efficiency.

This invention relates to a ground flight control and operation support system for a very low Earth orbit (UEO) satellite constellation. It incorporates predicted space environment information, satellite attitude change information, and orbit control information into the high-precision satellite orbit prediction process, effectively ensuring higher orbit prediction accuracy for the UEO constellation and maximizing the assistance to ground control stations during satellite transits in capturing and tracking the satellite. The satellite attitude and orbit control system simulation module performs real-time simulations of the impact of attitude control and orbit control on orbital maneuvers, thereby achieving higher orbit control accuracy and providing better assurance for the precise execution of missions by the UEO constellation. Automatic scheduling of Earth remote sensing and data transmission tasks, along with ground system fault identification and handling, and collision warning, provide more intelligent support for the operation of the UEO constellation.

A reusable launch vehicle comprising auxiliary tanks for non-ascent burns, such as orbital maneuver burns, reorientation burns, deorbit burns, or landing burns, and / or power generation. The auxiliary tanks may provide propellants for an orbital maneuvering system, an orbital reaction control system, a reentry reaction control system, and / or a power generation system. The tanks are configured to maintain cryogenic properties of auxiliary fuel and auxiliary oxidizer for the duration of the mission. An auxiliary fuel tank, a main fuel tank, an auxiliary oxidizer tank, and a main oxidizer tank may have various relative positions with respect to each other within the launch vehicle. The auxiliary fuel tank and auxiliary oxidizer tank may comprise insulation to further insulate the cryogenic fluids.

The application provides a multi-GEO satellite-based common view time transfer method, which comprises the following steps: selecting multiple GEO satellites in the visible range of two tracking stations, respectively using each GEO satellite to perform common view time transfer, and obtaining common view time transfer results of the GEO satellites; preprocessing the common view time transfer results of the GEO satellites to obtain preprocessed common view time transfer results; traversing the GEO satellites, analyzing the stability of the link of the current GEO satellite, and simultaneously referring to broadcast ephemeris to determine whether the current GEO satellite is in an orbit maneuvering state; if not, assigning a weight to the GEO satellite based on a static adaptive weight distribution method, or assigning a weight to the GEO satellite based on a dynamic adaptive weight distribution method; and based on the weights assigned to the GEO satellites, performing weighted average processing on the preprocessed common view time transfer results to finally obtain multi-GEO satellite common view time transfer results.

This invention discloses a trajectory planning method for adjusting the visibility time and uncertainty of a spacecraft, relating to the field of spacecraft technology. The method includes: establishing a spacecraft dynamics model considering perturbations and a geometric visibility constraint model for a situational awareness device; defining a vertex detection angle as an intermediate variable relating the spacecraft's visibility time and state uncertainty; solving for the vertex detection angle of the spacecraft relative to the situational awareness device based on latitudinal argument matching and Newton's iteration method; constructing an optimization problem with the spacecraft's pulse velocity increment as the optimization variable and maximizing or minimizing the square of the vertex detection angle as the optimization objective, and numerically solving it using a sequential quadratic programming algorithm to obtain the optimal pulse velocity increment; and performing spacecraft orbital maneuvers based on the optimal pulse velocity increment. This invention can establish the relationship between the orbital maneuver vector and the detected characteristic indicators, enabling rapid optimization of the spacecraft maneuver vector.

A solar sail includes a bus and a plurality of separate movable vanes coupled to the bus. Each movable vane includes a reflective surface for generating solar radiation pressure and propel the solar sail in space. Each vane may be movable relative to the bus in a fully deployed configuration such that an amount of thrust generated by solar radiation pressure on each vane is controllable.

The application discloses a design method of a high-mobility stealth satellite communication system and relates to the technical field of satellite communication. The application comprises the following steps: step a: system silence and state presetting: a control center makes a satellite of a space segment enter an electromagnetic silence state, a satellite platform starts surface thermal management and retracts a protruding structure, and an orbit maneuver with no fixed period is planned and executed. The application improves the concealment and survivability of the satellite platform in orbit. By making the satellite enter a silence and survival state and implementing thermal management and shape control, combined with the orbit maneuver with no fixed period, the physical and orbit characteristics of the satellite are difficult to be observed and predicted for a long time, traditional detection and tracking means based on a fixed shape and a periodic orbit are invalid, and exposure risk of space assets is controlled. Meanwhile, communication adopts a "zero handshake" burst access mechanism based on one-time instructions, and open signaling interaction in a link establishment process is completely eliminated.

Method for managing a satellite receiving CDM messages relating to a secondary object (2) that could collide with the satellite at a TCA date. At each orbit, an autonomous orbit control module establishes a safe current plan on board, defining station-keeping maneuvers (14, 15). A preliminary filtering of the received CDMs is performed on board based on geometric and / or temporal criteria. For each selected risky CDM, a collision risk level is estimated on board based on onboard navigation data propagated at the TCA date of said risky CDM. In the event of a confirmed risk, an ACA collision risk management module develops a new plan on board by removing or replacing at least one maneuver from the safe current plan with a maneuver (16, 17) that allows both station-keeping within a mission window (301-302) and avoidance. Figure for the abstract: Fig. 4

The invention relates to the technical field of satellite attitude control and orbital maneuver, in particular to a Brayton cycle-based satellite composite propulsion system and method. The composite propulsion system comprises a propellant storage unit which comprises a gaseous oxygen storage tank and a gaseous hydrogen storage tank; the turbocharging unit comprises an oxygen first-stage gas compressor and an oxygen second-stage gas compressor which are coaxially connected in series, a hydrogen first-stage gas compressor and a hydrogen second-stage gas compressor which are coaxially connected in series, and a turbine; the oxygen first-stage gas compressor, the oxygen second-stage gas compressor, the hydrogen first-stage gas compressor and the hydrogen second-stage gas compressor are coaxially connected with the turbine; an oxidant inlet of the fuel gas generator is communicated with an outlet of the oxygen secondary gas compressor, a fuel inlet of the fuel gas generator is communicated with an outlet of the hydrogen secondary gas compressor, and a fuel gas outlet of the fuel gas generator is communicated with an inlet of the turbine. A two-stage coaxial series-connection gas compressor structure is adopted, exhaust gas generated after a turbine does work is still kept to be higher than the pressure of a main thrust chamber, the exhaust gas can smoothly flow back to the main thrust chamber for afterburning, and full utilization of energy in a system is achieved.

The invention belongs to the technical field of satellite orbit maneuver detection, and particularly discloses an orbit maneuver detection method based on kinematic orbit epoch difference. According to the method, the kinematic orbit data is used for solving the orbit machine power through a non-difference method, so that the dependence on a prior dynamic model is eliminated, and a high-precision satellite-borne GNSS kinematic orbit product is directly utilized. The kinematic orbit is a pure geometric solution obtained through PPP (Precision Point Positioning), and is not smoothed through any dynamics, so that the position jump of the satellite during maneuvering is truly recorded. The method is simple and reliable, kinematics orbit data are easy to obtain, extra deviation products do not need to be provided externally, the algorithm is simple and easy to implement, and real-time application is convenient. In addition, the method also fills up the data blank during the shutdown period of the accelerometer, and can provide purer background field information for the inversion of the earth gravitational field.

The invention discloses a design method of a high-maneuverability stealth satellite communication system, and relates to the technical field of satellite communication. The method comprises the following steps: step a, system silent latent and state presetting: a control center enables a satellite in a space section to enter an electromagnetic silent state, a satellite platform starts surface thermal management and retracts a protruding structure, and meanwhile, orbital maneuver without a fixed period is planned and executed; according to the method, the concealment and survivability of the satellite platform existing in orbit are improved, the satellite enters the silent latent state, thermal management and appearance control are implemented, and in combination with orbit maneuver without a fixed period, the physical and orbit characteristics of the satellite are difficult to be stably observed and predicted for a long time; a traditional detection tracking means based on a fixed shape and a periodic track loses efficacy, the exposure risk of space assets is controlled, meanwhile, communication adopts a'zero handshake 'burst access mechanism based on a one-time instruction, and public signaling interaction in the link establishment process is thoroughly eliminated.

The present application provides a module-based automatically separable space propulsion system apparatus and a design method therefor. A storage and supply module has one end connected to an orbital maneuvering engine module via a first separation interface and another end connected to an attitude control engine module via a second separation interface. The storage and supply module provides propellants to the orbital maneuvering engine module and the attitude control engine module. The orbital maneuvering engine module provides thrust required for orbital maneuvering, and the attitude control engine module provides thrust required for attitude control. The first separation interface and the second separation interface are configured to allow separation among the storage and supply module, the orbital maneuvering engine module, and the attitude control engine module.

The application relates to a mega-constellation high-precision orbit rapid simulation method, device, equipment and medium. First, initial orbit information of a mega-constellation is acquired, rapid high-precision orbit simulation calculation is carried out through a preprocessing module, and orbit proxy model parameters are obtained; the orbit proxy model parameters are sent to a real-time calculation module for mega-constellation orbit simulation; when the real-time calculation module receives a new satellite control instruction, a calculation task of the new satellite control instruction is sent to a calculation pool module for simulation calculation, new orbit proxy model parameters are obtained; and the new orbit proxy model parameters are sent to the real-time calculation module for re-mega-constellation orbit simulation. Under the premise of ensuring high precision, the application realizes rapid simulation of tens of thousands of satellite orbits, can respond to constellation satellite orbit maneuvering control instructions in real time, and can rapidly adjust and calculate the satellite orbit.

A vehicle configured for in-space and atmospheric reentry operations includes a pressure-fed propulsion engine that uses cryogenic fuel and cryogenic oxidizer as propellants. In some embodiments, the vehicle includes a first conduit configured to provide a first propellant to the pressure-fed propulsion engine, a second conduit configured to provide a second propellant to a heat shield heat exchanger disposed relative to a heat shield wall defining an outer surface of the vehicle, and a pump located along the second conduit. In such embodiments, the first conduit is configured to provide the first propellant to the pressure-fed propulsion engine at a first mass flow rate, the second conduit and the pump are configured to provide the second propellant to the heat shield heat exchanger at a second mass flow rate, and the second mass flow rate is substantially less than the first mass flow rate. A related method is also disclosed.

This invention discloses a trajectory planning method for adjusting the visibility time and uncertainty of a spacecraft, relating to the field of spacecraft technology. The method includes: establishing a spacecraft dynamics model considering perturbations and a geometric visibility constraint model for a situational awareness device; defining a vertex-crossing detection angle as an intermediate variable relating the spacecraft's visibility time and state uncertainty; solving for the vertex-crossing detection angle of the spacecraft relative to the situational awareness device based on latitudinal argument matching and Newton's iteration method; constructing an optimization problem using the spacecraft's pulse velocity increment as the optimization variable and maximizing or minimizing the square of the vertex-crossing detection angle as the optimization objective, and numerically solving it using a sequential quadratic programming algorithm to obtain the optimal pulse velocity increment; and performing spacecraft orbital maneuvers based on the optimal pulse velocity increment. This invention can establish the relationship between the orbital maneuver vector and the detected characteristic indicators, enabling rapid optimization of the spacecraft maneuver vector.

This application relates to a method and system for obtaining joint maneuver parameters for geostationary orbit transfer and rendezvous. The method first obtains the initial parameters of the geostationary orbit transfer and rendezvous mission and determines the total number of orbital maneuvers, the total mission duration, and the rendezvous terminal aiming parameters. Then, it sequentially calls pre-constructed GTO transfer maneuver parameter acquisition problem models and GSO rendezvous maneuver parameter acquisition problem models for calculation. Next, it calls a pre-constructed joint maneuver parameter acquisition model for geostationary orbit transfer and rendezvous, considering the GTO transfer maneuver and the GSO rendezvous maneuver together. Finally, it uses a sequential quadratic programming algorithm to optimize and solve the joint maneuver parameter acquisition model for geostationary orbit transfer and rendezvous, obtaining the joint maneuver parameters. This allows for rapid rendezvous with the target spacecraft using the tracker's own thrust-limited orbital control engine, effectively shortening the orbital rendezvous mission time.

This invention discloses a reinforcement learning-based satellite pursuit and escape method and system based on orbital maneuver prediction. The method constructs a SAC policy network, designs a suitable loss function for backpropagation, and trains the initial policy of the model. A curiosity mechanism is introduced into the SAC algorithm, using the error between the prediction results and the actual values ​​from a fully connected network as a reward signal to encourage the agent to explore the action space. An input vector format and dataset for Transformer training are constructed, and the input vectors are positionally encoded upon input. A suitable loss function is designed for backpropagation of the Transformer. The Transformer's output is used as the input to the SAC algorithm, and the SAC observations are reconstructed, forming an integrated decision network. This invention solves the game problem of satellite pursuit and escape using pulse maneuvers.

An orbital maneuver apparatus includes a fluid supply arrangement, a main body, an actuation head, and a nozzle assembly. The main body includes an outer housing, an inner housing mounted in the outer housing such that the inner housing is arranged to rotate about a longitudinal direction of the main body, wherein the fluid supply tube extends in the inner housing of the main body and is stationary with respect to the outer housing. The actuation head is connected to the inner housing of the main body to rotate along with the inner housing. The nozzle assembly includes a first nozzle head and a second nozzle head which are in fluid communication with the fluid supply arrangement, and are both movably supported in the actuation cavity and are arranged to rotate correspondingly with the actuation head, and to further rotate about a transverse direction of the main body.

A thruster for orbital maneuvers including a divergent channel with axial axis of symmetry; a combustion chamber; a throat between the combustion chamber and the divergent channel; an injection plate facing the combustion chamber; first injection channels in fluid communication with the combustion chamber, each including an end portion open in the combustion chamber; at least one second injection channel in fluid communication with the combustion chamber, including an end portion open in the combustion chamber; wherein the end portions of the first injection channels are in an annular region extending along injection directions; wherein the end portion of the second injection channel is radially between the axis of symmetry and the end portions of the first injection channels; wherein the injection direction of the end portion of each first injection channel has axial and radial components directed toward the combustion chamber and the axis of symmetry, respectively.

The invention relates to the technical field of spaceflight measurement and control and space situation awareness, and discloses a satellite orbit maneuver detection method based on multichannel residual imaging, which comprises the following steps: acquiring an orbit element actual observation value, acquiring an orbit element theoretical value, calculating an orbit element residual sequence based on the orbit element actual observation value and the orbit element theoretical value, and calculating the orbit element maneuver according to the orbit element residual sequence. The method comprises the following steps of: extracting multiple feature channels, respectively carrying out Gramer angle field transformation on a residual sequence of each feature channel to obtain a multi-channel GAF image, constructing a classification model based on a deep neural network, and carrying out feature extraction and orbit maneuver judgment on the multi-channel GAF image. According to the method, the Gramb angle field transformation is applied to satellite orbit data processing for the first time, unequally spaced TLE time sequences are converted into regular two-dimensional images, the mature computer vision technology can be directly applied to maneuvering detection tasks, the time sequence correlation of the time sequences is reserved through GAF transformation, the time sequences are coded into space structure features of the images, and the space structure features of the images are obtained. And deep learning model extraction and learning are facilitated.

A reusable launch vehicle comprising auxiliary tanks for non-ascent burns, such as orbital maneuver burns, reorientation burns, deorbit burns, or landing burns, and / or power generation. The auxiliary tanks may provide propellants for an orbital maneuvering system, an orbital reaction control system, a reentry reaction control system, and / or a power generation system. The tanks are configured to maintain cryogenic properties of auxiliary fuel and auxiliary oxidizer for the duration of the mission. An auxiliary fuel tank, a main fuel tank, an auxiliary oxidizer tank, and a main oxidizer tank may have various relative positions with respect to each other within the launch vehicle. The auxiliary fuel tank and auxiliary oxidizer tank may comprise insulation to further insulate the cryogenic fluids.

The application relates to the technical field of constellation satellite maintenance regulation, and discloses a low-orbit Walker constellation configuration maintenance intelligent regulation system and method, which comprises the following steps: acquiring the variation trend of the in-plane phase angle and the inter-plane distance of a satellite in real time through a high-precision orbit prediction model; inputting the in-plane phase angle and the inter-plane distance into a pre-trained decision model to determine whether the parameters exceed a preset configuration maintenance threshold and to locate the unstable satellite; if the parameters exceed the limit, the model outputs the optimal orbit maneuvering parameters of the satellite based on the double-target constraints of minimum fuel consumption and shortest service interruption; fusing the phase maintenance requirements and the observation tasks on the satellite to generate an optimal control scheme, combining task planning and implementation to maintain the constellation configuration; and continuously monitoring the constellation configuration state of the satellite to maintain the stable operation of the constellation. The application realizes the automatic maintenance of the Walker constellation configuration, improves the control precision and efficiency.

The present invention discloses an autonomous system (100) for the removal of sub-10-centimeter orbital debris in Low Earth Orbit (LEO) and beyond. The system comprises an air tank module (102), a collector module (104), a deployable connector tunnel (112) forming a hermetically sealed debris transfer path, a laser alignment system (106) for axial coordination, an electromagnetic levitation system (108) for module separation and reattachment, ion propulsion systems (110) integrated on both modules for orbital maneuvering and positional control, and a centralized Al -based control unit (114) integrating an infrared imaging engine (114a) and a trajectory computation module (114b). The system also features a closed-loop gas recovery and re-pressurisation subsystem for sustainable multi-cycle operations. The system detects untracked debris, separating the modules to align on opposite sides of the debris, guiding, and propelling the debris centrally through the tunnel, capturing and compacting it internally, and reconnecting the modules to reset for the next cycle.