26 results about "Orbital operations" patented technology

The invention relates to the technical field of spacecraft autonomous time keeping, and discloses a satellite autonomous clock error estimation method, a reference establishment method, a device and a product. The satellite autonomous clock error estimation method comprises the following steps: constructing a laser communication link between satellites based on a laser communication list; any satellite carrying an atomic clock with the precision not lower than E-16 magnitude is determined as a reference satellite; calculating a target clock difference between a target satellite and the reference satellite based on a first clock difference between adjacent satellites in the laser communication link; the target satellite is any satellite except the reference satellite in the navigation satellite constellation system; and the adjacent satellites directly communicate with each other through a laser communication link. By adopting the method, a navigation constellation system running in a high orbit can realize space-based autonomous high-precision clock error estimation and clock synchronization in a global range.

Described herein is a method for wireless transmission of energy in space, wherein a constellation of transmitting satellites orbiting around a celestial body, such as the planet (P) Earth or the Moon (L), are equipped with a laser apparatus for sending a beam of coherent electromagnetic waves to an orbiting satellite or another receiving target, such as a mobile vehicle or a fixed base on said generic celestial body, in order to transmit energy thereto. The total power transmitted is divided among the orbiting satellites in the constellation, resulting in less contraindications than in the case of a single transmitting system with equivalent total power. The advantage of such division among a constellation of satellites mainly consists of shorter periods of solar eclipse, since sunlight is necessary for the operation of the transmission system. The receiving target or satellite converts the beam of electromagnetic waves into electric energy.

The invention discloses a method, a device and equipment for constructing a solar light pressure model of a low-orbit satellite, and relates to the technical field of satellite navigation and positioning. In the method, a method based on a carrier phase and an accelerometer is considered to calculate the total perturbative force acceleration of the low-orbit satellite, and the total perturbative force acceleration of the low-orbit satellite is calculated; environment state information of the low-orbit satellite is collected and combined with the total perturbation force acceleration to obtain the sunlight pressure acceleration, and based on the initial radiation coefficient used by the initial low-orbit satellite sunlight pressure model and combined with the sunlight pressure acceleration, the corrected radiation coefficient is obtained; and constructing a low-orbit satellite solar light pressure model of the first correction period by adopting the corrected radiation coefficient. In this way, the total acceleration of the perturbation force of the low-orbit satellite is calculated by using the satellite-borne carrier phase and the actually measured data collected by the accelerometer, and the corrected radiation coefficient is obtained based on the least square estimation to finely construct the solar light pressure model of the low-orbit satellite. And serious dependence of the light pressure model of the low-orbit satellite in orbit operation on prior information parameters is avoided.

The invention discloses a satellite orbit determination method and device, a storage medium and electronic equipment. The method comprises the following steps: acquiring an initial satellite position of a target communication satellite under the condition that the target communication satellite can receive a navigation signal, and determining a reference orbit of the target communication satellite based on spatio-temporal information of a navigation satellite; obtaining a target satellite position and a target non-conservative force model coefficient through orbit fitting; performing orbit integration to determine an integration orbit; comparing the integral orbit with the reference orbit to determine a position error sequence and process noise covariance information; and under the condition that the target communication satellite cannot receive the navigation signal, performing orbit filtering by using the inter-satellite link data, and determining a target operation orbit. The technical problem that a satellite cannot run according to a correct orbit due to the fact that a satellite constellation integrally deviates and rotates is solved.

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.

The invention presents a method for managing spatial data in a space and Earth surveillance or orbital operations security environment, the method being characterized by the following steps: Structuring the spatial data in the form of a directed acyclic graph (DAG); Associating each node with a creation date and traceability metadata allowing each node to be linked to the parent nodes from which it is derived; Including for each node numerical quality indicators and a confidentiality level specific to the data it contains; Transferring said metadata containing said indicators and levels from each node to the derived nodes; Managing and monitoring data transformations based on said quality and confidentiality metadata. Figure 2

A spacecraft (100) for functioning as an upper rocket stage (second stage or third stage) of a rocket during launch and as an orbital space station in orbit is disclosed The spacecraft (100) includes a payload section (Z3) have a plurality of modules adapted to facilitate orbital operation of the spacecraft (100). A central zone includes a compartmentalized tank module operationally coupled to the payload section (Z3). The central compartmentalized tank module having a plurality of compartments (7) adapted to store propellants during launch, transition into pressurized habitable compartments suitable for habitation and storage after venting and purging of residual propellants. A propulsion zone (Z1) is adapted to provide thrust during ascent and is adapted to detach from the central compartmentalized tank module upon achieving orbit.

The application discloses a low-orbit satellite solar radiation pressure model construction method, device and equipment, and relates to the technical field of satellite navigation and positioning. In the method, a carrier phase and an accelerometer are used to calculate the perturbation force total acceleration of a low-orbit satellite. The environmental state information of the low-orbit satellite is collected, and the solar radiation pressure acceleration is obtained in combination with the perturbation force total acceleration. Based on the initial radiation coefficient used in the initial low-orbit satellite solar radiation pressure model, the corrected radiation coefficient is obtained in combination with the solar radiation pressure acceleration. The low-orbit satellite solar radiation pressure model of the first correction period is constructed by using the corrected radiation coefficient. In this way, the measured data collected by the on-board carrier phase and the accelerometer is used to calculate the perturbation force total acceleration of the low-orbit satellite, and the corrected radiation coefficient is obtained based on the least square estimation to finely construct the low-orbit satellite solar radiation pressure model, thereby avoiding the serious dependence of the light pressure model of the low-orbit satellite on the prior information parameters in the orbit operation.

The application discloses a satellite orbit determination method and device, a storage medium and an electronic device. The method comprises the following steps: acquiring an initial satellite position of a target communication satellite in the case that the target communication satellite can receive a navigation signal, and determining a reference orbit of the target communication satellite based on space-time information of a navigation satellite; obtaining the target satellite position and a target non-conservative force model coefficient through orbit fitting; then performing orbit integration to determine an integrated orbit; comparing the integrated orbit and the reference orbit to determine a position error sequence and process noise covariance information; and performing orbit filtering using intersatellite link data in the case that the target communication satellite cannot receive the navigation signal, so as to determine a target operation orbit. The application solves the technical problem that the satellite cannot operate according to a correct orbit due to overall offset rotation of a satellite constellation.

The application provides a satellite orbit prediction method and device based on orbit data of a spherical satellite. The method comprises determining a target atmospheric density model according to a target space environment condition of a satellite to be predicted, wherein the target atmospheric density model is obtained by correcting a preset atmospheric density model according to orbit data and a dynamic model of the spherical satellite in the target space environment condition; identifying an atmospheric drag coefficient of the satellite to be predicted based on the target atmospheric density model; and performing orbit prediction on the satellite to be predicted according to the target atmospheric density model and the atmospheric drag coefficient. According to the embodiment of the application, the spherical satellite maintains a stable and high-precision surface value ratio and atmospheric drag coefficient when operating in an orbit, the orbit data of the spherical satellite can be used to better achieve high-precision correction of the preset atmospheric density model, and the atmospheric drag coefficient of the satellite to be predicted can be more accurately identified based on the corrected atmospheric density model, thereby improving the orbit prediction accuracy of the satellite to be predicted.

The invention relates to the field of satellite communication, in particular to a high and low orbit satellite communication method, device and equipment and a storage medium. The method comprises the following steps: identifying orbit moving tracks of high and low orbit satellites, carrying out geometric topology change modeling, and constructing a satellite topology change model; performing effective communication orbit calculation based on the satellite topology change model to obtain a communicable orbit area; performing multi-index communication channel evaluation on the communicable track area to generate real-time channel state parameters; solving an optimal communication position based on the real-time channel state parameter to obtain an optimal communication position node coordinate; and performing synchronous communication link based on the optimal communication position node coordinates, and performing communication transmission protocol stack optimization to complete high-orbit and low-orbit satellite communication operation. According to the method, the link is established in the optimal communication path, so that the communication delay is reduced, and the link stability is improved.

For maneuvering a spacecraft, a method calculates a virtual point that represents a plurality of spacecraft orbiting in a spacecraft formation. The method iteratively calculates a desired trajectory relative to the virtual point for a given spacecraft of the plurality of spacecraft using a relative orbital elements (ROE) state x of the virtual point and the given spacecraft. Model predictive control maintains the given spacecraft within an ROE constraint for a given time interval subject to a fuel consumption. In response to determining a drift trajectory of the given spacecraft is outside of a trajectory tolerance of the desired trajectory, maneuvering the given spacecraft to within the trajectory tolerance using a maneuver trajectory determined using the model predictive control.

This application relates to the field of satellite orbit control technology, and particularly to a semi-analytical attitude adjustment and unloading method and apparatus for a Halo orbit satellite at the Earth-Moon L2 point, to improve the computational efficiency and accuracy of the satellite's active attitude unloading. It includes: determining the initial velocity increment required for orbit maintenance based on basic information about the satellite's current orbital operation; determining the desired velocity increment for orbit maintenance in the current orbit maintenance direction based on the initial velocity increment; representing the remaining velocity increment for attitude adjustment and unloading in a reference coordinate system using azimuth and elevation angles; optimizing the azimuth and elevation angles in the reference coordinate system with the remaining velocity increment for attitude adjustment and unloading as the optimization objective; determining a reference attitude adjustment and unloading attitude based on the optimized azimuth and elevation angles; obtaining the actual orbit maintenance velocity increment in the direction corresponding to the reference attitude adjustment and unloading attitude; correcting the reference attitude adjustment and unloading attitude based on the actual orbit maintenance velocity increment; and determining the satellite's target unloading attitude.

The embodiment of the application discloses a two-degree-of-freedom discontinuous rotation control method and system for a solar wing, the discontinuous rotation control method comprises the following steps: obtaining the angle between a sun vector and the normal of a solar wing along the direction of orbit operation and the angle between the sun vector and the normal of the solar wing in the direction perpendicular to the orbit plane, so as to obtain the angle between the sun vector and the normal of the solar wing in a first coordinate projection plane and a second coordinate projection plane, and further obtain the driving target instruction of a first control shaft and a second control shaft; determining the position of the satellite, if the satellite is in a sunlit area, driving the first control shaft to make the sun vector coincide with the normal of the solar wing in the first coordinate projection plane, then driving the second control shaft to adjust the position angle of the solar wing relative to the orbit plane, so that the normal of the solar wing coincides with the sun vector in the second coordinate projection plane; if the satellite is in a shadow area, driving the first control shaft to execute a rotation instruction.

A spacecraft (100) for functioning as an upper rocket stage (second stage or third stage) of a rocket during launch and as an orbital space station in orbit is disclosed The spacecraft (100) includes a payload section (Z3) have a plurality of modules adapted to facilitate orbital operation of the spacecraft (100). A central zone includes a compartmentalized tank module operationally coupled to the payload section (Z3). The central compartmentalized tank module having a plurality of compartments (7) adapted to store propellants during launch, transition into pressurized habitable compartments suitable for habitation and storage after venting and purging of residual propellants. A propulsion zone (Z1) is adapted to provide thrust during ascent and is adapted to detach from the central compartmentalized tank module upon achieving orbit.

The invention discloses a transfer orbit design method in on-orbit service of a synchronous orbit satellite. The problem of how to effectively image a high-orbit target satellite with low fuel consumption is solved. Belongs to the technical field of space. Comprising the following steps: determining total transfer time according to different initial states of a service satellite and a target satellite; constructing a transfer orbit mathematical model based on the total transfer time to obtain the running time of the service satellite on the initial orbit, the transfer orbit and the approaching orbit in the transfer process and the speed increment for completing the Horman transfer; taking a relative position vector included angle and an illumination angle of two stars at the end of intersection as an objective function, taking total transfer time as a constraint condition, taking the operation time of a service star in an initial orbit and an approaching orbit as an optimization variable, optimizing the objective function by using a genetic algorithm based on an HPOP model, and correcting parameters of a transfer orbit mathematical model to obtain a transfer orbit mathematical model; obtaining a transfer orbit mathematical optimization model; according to the method, effective imaging of the high-orbit target with low burnup is realized through the mathematical optimization model of the transfer orbit.

The present invention relates to a fastening system for securing an optical element, an optical assembly, a method for securing an optical element, and a method for assembling and aligning an optical assembly. Specifically, the invention relates to a fastening system for securing optical elements for a camera intended for being placed in a satellite orbiting in space.

According to the device for simulating the potential state of the spacecraft structure, the potential state of the spacecraft structure running in different orbits and in a complex space environment is simulated through an electron injection mode and an electron extraction mode, and the device is an efficient and controllable space environment simulation tool; the device has the characteristics of adjustable electron flow, simple equipment, high integration level, compact structure and wide application range, is an important bridge for connecting space environment theory, material science and aerospace engineering practice, and can be applied to potential simulation experiments of different spacecrafts; the platform is an indispensable key ground verification platform for ensuring high-reliability and long-life operation of a spacecraft in orbit.

Described are techniques for determining an impact threat of an orbiting object by a crowdsource satellite network. The techniques include capturing, by a satellite included in a crowdsource satellite network, information associated with an orbiting object detected by the satellite. The techniques further include providing, by the satellite, the information associated with the orbiting object to the crowdsource satellite network. The techniques further include collaborating by the satellite with other satellites in the crowdsource satellite network to track the orbiting object, generate an orbital model of the orbiting object, and to analyze a predicted path of the orbiting object represented by the orbital model to determine an impact threat associated with the predicted path of the orbiting object.

The invention discloses a rolling attitude maneuvering sailboard sun-pointing method for an aircraft running on an inclined low orbit, and belongs to the technical field of autonomous sun-pointing design of sailboards of on-orbit aircrafts. The method comprises the following steps: firstly, calculating a sun vector under an inertial system, and then calculating a solar altitude angle, a solar azimuth angle and a solar pitch angle under an aircraft orbital system; designing an optimal target attitude angle of the aircraft based on constraint conditions; finally, after the attitude of the aircraft is adjusted in place, the sailboard rotates, and the rotation angle and the rotation speed of the sailboard are designed. According to the method, the corresponding optimal target attitude angle and the sailboard rotation mode are designed based on different solar altitude angles, long-term on-orbit sailboard sun alignment is achieved, and energy supplement is guaranteed.

The invention provides a giant satellite base simulation system and method based on OMNeT + +, and can solve the problem of insufficient efficiency and precision in giant satellite base simulation software development in related technologies. The system comprises a path planning module which is used for constructing a dynamic weight model by adopting a dynamic weight algorithm to obtain a path dynamic weight and generating an optimal planning path of a satellite according to the path dynamic weight; the working mode adjusting module is used for monitoring the running state of the system in real time, confirming a target working mode according to the label information of each module in the system, and sending an adjusting signal for adjusting the target working mode to each module in the corresponding system; the protocol simulation module is used for processing various protocol messages in the handleMessage function and then transmitting simulation result data to the simulation function implementation module; and the simulation function realization module is used for completing orbit operation simulation of the satellite according to the optimal planning path, the adjustment signal and the simulation result data.

The application provides a long-term residence method and system of an earth-pole-suspension spacecraft, and relates to the technical field of spaceflight, and comprises the following steps: S1, a graphene film sail surface is arranged for the earth-pole-suspension spacecraft, the graphene film sail surface is used for interacting with the sunlight, and a suspension centripetal force is calculated; S2, a control system is arranged, the included angle relationship between the graphene film sail surface and the sunlight is controlled according to the suspension centripetal force, the requirement of a suspension orbit is satisfied in real time, a spiral suspension orbit operation is formed, the earth revolves around the sun, one-year return is formed, and the earth-pole-suspension orbit operation of the next cycle is continued. The application can effectively improve the comprehensive detection performance and observation efficiency of the polar spacecraft.

The application provides a super constellation simulation system and method based on OMNeT++, which can solve the problems of insufficient efficiency and precision in the development of super constellation simulation software in the related art. The system comprises: a path planning module, configured to adopt a dynamic weight algorithm, construct a dynamic weight model to obtain path dynamic weights, and generate an optimal planning path of a satellite according to the path dynamic weights; a working mode adjustment module, configured to monitor the system operation state in real time, confirm a target working mode according to label information of each module in the system, and send an adjustment signal for adjusting the target working mode to each module in the corresponding system; a protocol simulation module, configured to transmit simulation result data to a simulation function implementation module after processing various protocol messages in a handleMessage function; and a simulation function implementation module, configured to complete orbit operation simulation of the satellite according to the optimal planning path, the adjustment signal and the simulation result data.

The invention discloses an intelligent breathable eye socket post-operation eye protection cover, and relates to the technical field of medical instruments, the intelligent breathable eye socket post-operation eye protection cover comprises a mirror frame, vent holes are formed in the side face of the mirror frame, a baffle is arranged in the mirror frame, the baffle slides in the mirror frame through a sliding assembly to open or block the vent holes, and a temperature and humidity monitoring module is arranged in the mirror frame. The temperature and humidity monitoring module controls the sliding assembly in a feedback mode, and a detachable flexible cushion layer is arranged on the inner side of the mirror frame in a circle. The air permeability is adjusted through mutual cooperation of the baffle and the air holes, the slippage degree of the baffle is fed back and adjusted through the temperature and humidity monitoring module, the air permeability can be automatically adjusted according to the temperature and humidity changes of eye skin, and eyes are kept dry and comfortable. The temperature and humidity monitoring module is integrated around the eyes, the humidity and temperature data of the eyes are monitored in real time, the data are transmitted to the mobile equipment through the wireless communication module, abnormal conditions are reminded in time, and monitoring and management of postoperative recovery are facilitated.

This invention is embodied in a system and device for transporting assets (small satellites, commodities, or other payloads) in space. The preferred system comprises a scaffold of relay stations strategically orbiting one or more celestial bodies. The orbits are complimentary to each other so that the relay stations can transfer mass from one station to another. In this way, the system can support different logistical networks connecting the space between orbits such as Low Earth orbit (LEO), geostationary orbit (GEO), and Low Lunar orbit (LLO), and even between Earth and Mars.

The invention provides a thermal analysis and temperature prediction method and system for a multi-orbit cycle aerospace satellite, and the method comprises the steps: building a satellite three-dimensional physical model through finite element simulation software, setting the thermophysical parameters of all part materials, configuring a heat source and a contact thermal resistor in a solid heat transfer module, and carrying out the thermal analysis and temperature prediction. A space thermal environment and a satellite orbit attitude are defined in an orbit thermal load module, dynamic periodic control of heat source power is realized by utilizing an event module, and a step-by-step solving strategy is adopted: firstly, single-orbit periodic thermal load distribution is calculated, and then the single-orbit periodic thermal load distribution is periodically and repeatedly loaded in a plurality of orbit periods, so that the dynamic periodic thermal load control is realized. And finally, the temperature field change of the satellite in multi-orbit operation is obtained. The method effectively solves the problem that a traditional analysis method is difficult to give consideration to the multi-orbit periodic temperature accumulation effect and the dynamic thermal load change, significantly improves the prediction precision and calculation efficiency of the satellite thermal analysis, and provides a reliable basis for the thermal design of a spacecraft.